February 2006 

NASA/TM-2006-214273 

Aerocapture Systems Analysis for a Titan Mission 

Mary K. Lockwood, Eric M. Queen, David W. Way, Richard W. Powell, Karl Edquist, Brett W. Starr, 

Brian R. Hollis, E. Vincent Zoby, and Glenn A. Hrinda 

NASA Langley Research Center, Hampton Virginia 

Robert W. Bailey, Jeffery L. Hall, Thomas R. Spilker, Muriel Noca, and Robert Haw 

Jet Propulsion Laboratory, Pasadena, California 

Carl G. Justus and Aleta Duvall 

Computer Sciences Corporation, Huntsville, Alabama 

Dale L. Johnson 

NASA Marshall Space Flight Center, Marshall Space Flight Center, Alabama 

James Masciarelli 

NASA Johnson Space Center, Houston, Texas 

Naruhisa Takashima 

AMA Inc., Hampton, Virginia 

Kenneth Sutton 

National Institute of Aerospace, Hampton, Virginia 

Joe Olejniczak and Bernard Laub 

NASA Ames Research Center, Moffett Field, California 

Michael J. Wright and Dinesh Prabhu 

ELORET Corporation, Sunnyvale, California 

R. Eric Dyke 

Swales Aerospace, Hampton, Virginia 
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February 2006 

NASA/TM-2006-214273 

Aerocapture Systems Analysis for a Titan Mission 

Mary K. Lockwood, Eric M. Queen, David W. Way, Richard W. Powell, Karl Edquist, Brett W. Starr, 

Brian R. Hollis, E. Vincent Zoby, and Glenn A. Hrinda 

NASA Langley Research Center, Hampton Virginia 

Robert W. Bailey, Jeffery L. Hall, Thomas R. Spilker, Muriel Noca, and Robert Haw 

Jet Propulsion Laboratory, Pasadena, California 

Carl G. Justus and Aleta Duvall 

Computer Sciences Corporation, Huntsville, Alabama 

Dale L. Johnson 

NASA Marshall Space Flight Center, Marshall Space Flight Center, Alabama 

James Masciarelli 

NASA Johnson Space Center, Houston, Texas 

Naruhisa Takashima 

AMA Inc., Hampton, Virginia 

Kenneth Sutton 

National Institute of Aerospace, Hampton, Virginia 

Joe Olejniczak and Bernard Laub 

NASA Ames Research Center, Moffett Field, California 

Michael J. Wright and Dinesh Prabhu 

ELORET Corporation, Sunnyvale, California

R. Eric Dyke 

Swales Aerospace, Hampton, Virginia
Available from: 

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7121 Standard Drive 5285 Port Royal Road 

Hanover, MD 21076-1320 Springfield, VA 22161-2171 

(301) 621-0390 (703) 605-6000 

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not 

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the 

National Aeronautics and Space Administration. 
 

iii

TABLE OF CONTENTS 

Titan Aerocapture Systems Analysis ............................................. 1 

Titan Aerocapture Mission and Spacecraft Design Overview ............................................. 13 

Titan Explorer Mission Trades from the Perspective of Aerocapture ............................................. 24 

Approach Navigation for a Titan Aerocapture Orbiter ............................................. 35 

Engineering-Level Model Atmospheres for Titan and Neptune ............................................. 47 

Guidance Algorithms for Aerocapture at Titan ............................................. 53 

Aerocapture Simulation and Performance for the Titan Explorer 

Mission 

............................................. 61 

Preliminary Aerothermodynamics of Titan Aerocapture Aeroshell ............................................. 72 

An Analysis of the Radiative Heating Environment for Aerocapture at 

Titan 

............................................. 83 

Thermal Protection Concepts and Issues for Aerocapture at Titan ............................................. 92 

Structural Design of the Titan Aerocapture Mission ............................................. 102 

Planetary Probe Mass Estimation Tool Development and its 

Application to Titan Aerocapture 

............................................. 110 
1

TITAN AEROCAPTURE SYSTEMS ANALYSIS 

Mary Kae Lockwood 

NASA Langley Research Center

INTRODUCTION 

Performance projections for aerocapture show a 

vehicle mass savings of between 40 and 80%, 

dependent on destination, for an aerocapture vehicle 

compared to an all-propulsive chemical vehicle. In 

addition aerocapture is applicable to multiple 

planetary exploration destinations of interest to the 

NASA Office of Space Science. These results led to 

the identification of aerocapture as one of the top 

three propulsion technologies for solar system 

exploration missions during the 2001 NASA InSpace Propulsion Program (ISP) technology 

prioritization effort, led by Marshall Space Flight 

Center, to rank current ISP propulsion technologies. 

An additional finding was that aerocapture needed a 

better system definition and that supporting 

technology gaps needed to be identified. 

An aerocapture systems analysis effort was kicked 

off in late February and completed in September 

2002. The focus of the effort was on aerocapture at 

Titan with a rigid aeroshell system. Titan was 

selected as the initial destination for the study due to 

potential interest in a follow-on mission to 

Cassini/Huygens. The systems analysis is being 

completed by a multi-center NASA team including 

scientists and engineers from Ames Research Center, 

the Jet Propulsion Laboratory, Johnson Space Center, 

Langley Research Center, and Marshall Space Flight 

Center, led by Langley Research Center. Continued 

aerocapture systems analysis work is in progress with 

a Neptune aerocapture systems analysis study. 

Neptune is representative of the gas giant planets. 

Additional destinations and further work will be 

defined based on NASA Office of Space Science 

roadmap updates and ISP technology development. 

Plans in FY04 include Mars and Venus. 

SYMBOLS/NOMENCLATURE 

A Area (m2

) 

AFE Aeroassist Flight Experiment 

αtrim Trim Angle of Attack 

BOC Beginning of Cassini 

CA Axial Force Coefficient 

CBE Current Best Estimate 

CD Coefficient of Drag 

CFD Computational Fluid Dynamics 

CG, cg Center of Gravity 

CL Coefficient of Lift 

CN Normal Force Coefficient 

D Drag 

EGA Earth Gravity Assist 

EOC End of Cassini 

HGA High Gain Antennae 

L Lift 

L/D Lift-to-Drag ratio 

M/CDA Ballistic Coefficient (kg/m2

) 

SEP Solar Electric Propulsion 

TPS Thermal Protection System 

VGA Venus Gravity Assist 

BACKGROUND 

An aerocapture flight profile schematic is shown in 

Figure 1. The vehicle approaches the planet/moon 

from a hyperbolic approach trajectory, shown at point 

1, designed to achieve state conditions including 

Figure 1. Aerocapture flight schematic. 
2

flight path angle at atmospheric interface, point 2, 

within a predetermined range. Bank angle 

modulation, rotation of the lift vector about the 

velocity vector is initiated by the guidance at point 3. 

The drag on the vehicle as it passes through the 

atmosphere provides the delta V required to capture 

the vehicle into the desired orbit. The amount of delta 

V imparted to the vehicle is controlled by the onboard guidance by modulating bank angle, i.e. the 

direction of the vehicle’s lift vector. A command of 

lift up during an atmospheric pass results in 

increasing altitudes, nominally decreasing 

atmospheric density, reduced drag and reduced delta 

V imparted. A command of lift down results in 

decreasing altitudes, nominally increasing atmospheric density, increased drag and increased delta V 

imparted. Bank angle modulation is commanded 

throughout the atmospheric pass from point 3 to point 

5. By point 5, where the influence of aerodynamic 

forces is no longer significant, the energy depleted 

from the initial hyperbolic trajectory is that required 

to capture the vehicle into the desired orbit. At 

apoapsis, point 7, a small delta V burn is performed 

to raise the periapsis. 

AEROCAPTURE CORRIDOR 

The aerocapture theoretical corridor is bounded by 

the full lift up trajectory and the full lift down 

trajectory for a nominal atmosphere and vehicle 

aerodynamics. (The theoretical corridor width is 

defined by the difference between the entry flight 

path angle corresponding to a full lift down trajectory 

and the entry flight path angle corresponding to a full 

lift up trajectory for a nominal atmosphere and 

aerodynamics.) If the vehicle enters the atmosphere 

at a flight path angle steeper than defined by the full 

lift up trajectory the vehicle lands. If the vehicle 

enters the atmosphere at a flight path angle shallower 

than that defined by the full down trajectory the 

vehicle is not captured. Figure 2 illustrates the 

theoretical corridor and the effects of navigation, 

atmosphere and aerodynamic uncertainties on the 

theoretical corridor. The plot shows entry flight path 

angle for a full lift up and full lift down trajectory as 

a function of the atmosphere variable Fminmax, 

where Fminmax=0 is the nominal atmosphere, 

Fminmax=-1 is the lowest density atmosphere, 

Fminmax =+1 is the maximum density atmosphere. 

The plot is shown for a given vehicle, entry velocity, 

and target orbit. To first order, the aerocapture 

corridor required to accommodate atmospheric 

dispersions, navigation errors (delivery flight path 

angle), and aerodynamic uncertainties can be root 

sum squared to determine the total corridor required. 

If the theoretical corridor is significantly greater than 

the corridor width required, then the vehicle control 

authority is adequate and the aerocapture is robust. 

The approach is only an estimate. For example 

factors such as high frequency atmospheric density 

perturbations are not included in this approach and 

can affect the results. Monte Carlo simulation 

analyses must be completed to assess feasibility and 

robustness. 

Concerns expressed regarding the risk of aerocapture 

have largely been in three areas. 1) Atmosphere 

variability and uncertainty; 2) Approach navigation 

delivery errors; 3) Aerodynamics utilized to control 

capture. To address these concerns, the following 

approach can be taken in the design of the 

aerocapture system. 

• Based on available atmospheric measurements, 

quantify and model the physical range of 

atmosphere variability and uncertainty. 

• Quantify the approach navigation delivery 

errors, and incorporate navigation systems into 

the vehicle design to reduce errors. 

• Quantify the aerodynamic uncertainties 

including margin. 

• Select a vehicle L/D to provide adequate 

control authority. 

• Provide adequate vehicle control 

responsiveness to accommodate perturbations, 

including atmospheric perturbations. 

• Develop a robust guidance. 

• Evaluate aerocapture robustness through Monte 

Carlo simulation incorporating all variability, 

uncertainty, errors and dispersions. 

• Design the aerocapture system to provide 

margin above 3-σ success, in particular for first 

time flights and for high value payloads. 

Figure 2. Aerocapture flight corridor, effect of 

uncertainties and dispersions. 
3

PREVIOUS AEROCAPTURE ASSUMPTIONS 

Previous aerocapture mission analysis work has 

demonstrated significant benefit of aerocapture for 

outer planet missions. However, due to the broad 

nature of the work, the analysis was low fidelity. 

Many assumptions were made, and it was noted that 

higher fidelity analyses would be needed to address 

the assumptions. 

Assumptions made in previous aerocapture mission 

analyses were that the available aerocapture corridor 

width at a given destination was adequate to 

accommodate all of the dispersions and uncertainties, 

such as atmosphere uncertainties, navigation 

dispersions, and aerodynamic uncertainties. And that 

these uncertainties and dispersions could be 

quantified. It was assumed that guidance algorithms 

could be designed to successfully capture the vehicle 

over the range of uncertainties/dispersions. For Titan, 

the aerocapture subsystem mass, including structure 

and TPS, was estimated to be 27% of entry mass. It 

was assumed that aeroheating environments for 

aerocapture could be quantified and accommodated 

by TPS systems within the mass allocation. With 

entry velocities at Titan of 6-10 km/sec and the Titan 

atmosphere being predominantly Nitrogen, 

aeroheating rates were expected to be modest and 

convective heating dominated. As a result, TPS 

materials such as SLA 561 were expected to be 

applicable. All of the earlier studies assumed that the 

spacecraft could be packaged in the aeroshell; that 

volume is available, subsystems can be located to 

meet center of gravity (cg) restrictions, thermal and 

communication issues can be solved. It was assumed 

that the selected science orbits are feasible and that 

technology, including analysis tools would be ready 

in time to meet a project schedule. It was also 

estimated that the benefit of aerocapture at Titan 

compared to an all-propulsive capture provides a 

mass savings of 66%. 

All of the assumptions made were to be addressed in 

the systems analysis study described here. 

OBJECTIVES 

The objectives for the Titan aerocapture systems 

analysis were therefore to provide higher fidelity 

analyses for validation and update to aerocapture 

assumptions made in mission studies, including 

performance, environments, mass properties, etc. The 

results of the analysis were to be provided to 

scientists, mission planners, technology planners, 

technologists and future mission managers. The 

feasibility, benefit and risk of aeroshell aerocapture 

system and technologies for Titan destination were to 

be defined. Technology gaps were to be identified 

and performance goals of key technologies defined. 

APPROACH 

A multi-center aerocapture systems analysis team 

was formed, including NASA engineers and 

scientists from Ames Research Center, the Jet 

Propulsion Laboratory, Johnson Space Center, 

Langley Research Center, and Marshall Space Flight 

Center, led by Langley Research Center. The team 

kicked off the study in late February and completed 

the work in September 2002. 

The analysis included top level sensitivity studies to 

identify a reference concept for higher fidelity 

analysis and to provide sensitivities through a broader 

range of possible aerocapture mission scenarios. The 

reference concept was to provide a higher fidelity 

reference to address the previous assumptions noted 

above, to provide a reference for higher fidelity 

component level trades, and to provide a benchmark 

to the top-level sensitivities. 

The mission objectives and initial spacecraft design 

for the reference concept was based on JPL’s TeamX 

study1

 of the Titan Explorer mission. 

TITAN AEROCAPTURE REFERENCE 

CONCEPT 

The level one objectives for the Reference Concept 

were defined based on a modified set of those used in 

the TeamX study. The Titan Explorer consisted of an 

Orbiter and a Lander, each delivered to Titan. The 

Orbiter delivers the Lander to its Titan entry 

trajectory. The Lander performs a direct entry. The 

Orbiter aerocaptures into a near polar orbit about 

Titan. 

One of the goals is to minimize trip time to Titan. The 

science mission is to be three years, with three years 

of Orbiter operations at Titan. In addition, the Orbiter 

serves as a relay for the Orbiter for the first year. The 

Orbiter science instruments include a Microwave 

spectrometer, SAR, Multispectral imager, USO as 

described in more detail in Ref 4. The launch date 

selected is 2010. This requires technologies to be at a 

technology readiness level (TRL) of six by 2006. 

Other launch dates are also considered. A SEP 

propulsion module is selected for the Reference 

Concept with comparisons made to the chemical 

propulsion module. Given the relatively near-term 

launch date, as much heritage hardware as possible is 

utilized. The mission is defined as a Class A mission 

with a fully redundant design. The team was funded 

to study the aerocapture Orbiter only. The Lander had 

to be treated as a “black box” such that most of the 
4

TeamX analysis for the Lander was taken as is. The 

Lander mass allocation is 400 kg. 

CONCEPT SELECTION 

Selection of a Reference Concept is a balance 

between providing reduced trip time to the 

destination, adequate delivered mass, adequate 

aerocapture flight corridor width to provide a robust 

system, all while meeting the science objectives. 

Mission analyses were conducted for the Titan 

aerocapture mission with various launch vehicles, 

gravity assist options, chemical propulsion vs. solar 

electric propulsion (SEP), and various launch dates as 

described in Reference 3. Figure 3 shows the 

delivered mass vs. flight time for various launch 

dates for a Venus Gravity Assist (VGA) and a SEP 

propulsion module. Note that for flight times greater 

than six years, the maximum mass that can be 

delivered remains nearly constant. Delivered mass 

decreases significantly as trip time is reduced from 

six to five years or less. Figure 4 shows the inertial 

entry velocity for the same missions as Figure 2. Six 

to eight year trip times result in entry velocities of 

less than ~7 km/sec for most launch opportunities. 

Also, entry velocity increases rapidly for trip times 

less than five to six years. 

Since convective heating increases approximately 

with the cube of entry velocity as shown in Figure 4, 

trip times greater than six years were expected to 

reduce the mass of the required thermal protection 

system allowing more of the delivered mass to be 

allocated to meet science requirements. 

Selection of a 6.5 km/sec entry velocity with an 

approximately 6 year trip time, dependent on 

opportunity, resulted in a balance between trip time 

and mass. From Figure 5, the stagnation point 

convective aeroheating on a one meter nose radius at 

6.5 km/sec is approximately 40 W/cm2

. SLA 561 was 

expected to be adequate for the aeroheating 

environment. 

Note that due to the Titan orbit about Saturn, entry 

velocity can be increased for given flight time 

without loss in delivery mass capability. The 

velocities in Figure 4 are each minimum inertial entry 

velocity. 

To determine the theoretical corridor width leading to 

determination of the vehicle lift to drag ratio required, full lift up and full lift down aerocapture trajectories were developed over the range of entry velocities of interest, and for a range of lift to drag ratio 

vehicles. Initial analysis had shown that ballistic 

Figure 3. Delivered mass vs. mission flight time. 

Ref 3. 

Figure 4. Inertial entry velocity vs. mission flight time. 

Ref 3. 

Figure 5. Convective stagnation point heat rate vs. 

entry velocity. Ref 7. 
5

coefficient and target apoapsis were secondary 

variables in the determination of corridor width. 

From Figure 6, for a vehicle with L/D = 0.25, a 3.5 

deg theoretical corridor width with 6.5 km/sec entry 

velocity is achieved. (With a 10 km/sec entry and the 

same L/D, 4.7 deg theoretical corridor width is 

achieved.) The 3.5 deg corridor width was expected 

to be more than adequate to accommodate 3-σ

navigation delivery errors, atmosphere dispersions 

and aerodynamic uncertainties with 99.7% or greater 

success. This allowed a high heritage low L/D sphere 

cone configuration to be selected. If increased corridor 

were required, mid L/D configurations are viable 

alternatives. 

Figure 6. Aerocapture corridor width as a 

function of vehicle lift to drag ratio. Ref 7. 

NAVIGATION 

Several navigation approaches were analyzed for 

Titan as described in Reference 2. The approach 

selected for the Reference Concept is based on the 

Cassini optical navigation camera capability, ∆DOR 

(utilized on Odyssey), and Dopler and Range 

combined. The 3-σ flight path angle dispersions, 

based on a Beginning of Cassini (BOC) Titan 

ephemeris is ±1.42o

. Significant improvements in 

knowledge of the location of Saturn, Titan will result 

from Cassini mission. 3-σ sets of entry states, BOC 

and EOC, were utilized in the Monte Carlo analyses 

described below. This detailed navigation analysis 

was required to assess feasibility and robustness of 

Titan aerocapture. Further reduction in the flight path 

angle dispersions can be achieved with the resolution 

of the Mars Reconnaissance Orbiter camera, as noted 

in Reference 2. 

ATMOSPHERE MODELING 

TitanGRAM, discussed in detail in Reference 5, 

includes a model of measurement uncertainties, 

residual uncertainties (turbulence, waves, etc); 

variation with latitude, altitude, time of day, season. 

This model fidelity is required to assess mission 

feasibility and robustness. Figure 7 shows the Titan 

atmospheric density with altitude. 

The arrival date of the current study results in the 

maximum variation of density with latitude. Since the 

science orbit is near polar, aerocapture occurs over a 

wide range of latitudes. Figure 8 represents a 

simulation of the expected variability of Fminmax 

with latitude. 

Figures 7 and 8 show the mean variation of the Titan 

atmosphere. Also included in TitanGRAM are the 

atmosphere perturbations. Both the mean variations 

and perturbations are utilized in the Monte Carlo 

analysis described below. 

Cassini-Huygens data will reduce measurement 

uncertainty as discussed further below. 

Fminmax=1

Fminmax=0

Fminmax=-1

Figure 7. TitanGRAM atmospheric density as a 

function of altitude. 

Figure 8. Simulated variation of Fminmax 

TitanGRAM parameter with Titan latitude. 
6

AERODYNAMICS 

Aerodynamic uncertainties included in the Monte 

Carlo are based on LAURA CFD for blunt bodies at 

hypersonic velocities.7

 The uncertainties include ±3% 

uncertainty in CA; ±5% uncertainty in CN; and a trim 

angle of attack (αtrim) uncertainty of ±2 deg used to 

represent uncertainty in Cm. The CG uncertainty 

used is ±0.0318 m in axial, Zcg and ±0.0069 m in 

radial Xcg. Figure 9 (Ref 7) shows the effect of these 

uncertainties on vehicle lift to drag ratio. 

GUIDANCE 

HYPAS guidance (ref 6) was chosen for the Titan 

aerocapture systems analysis. (Other guidance 

algorithms are planned for future consideration.) 

HYPAS was originally developed for the Aeroassist 

Flight Experiment (AFE). However the current 

version includes several improvements since AFE. 

HYPAS utilizes vehicle lift and bank angle control 

through the atmosphere to target the desired exit orbit 

apoapsis and inclination (or plane). It is an 

analytically derived algorithm based on deceleration 

due to drag and altitude rate error feedback. This 

analytic, non-iterative, on-the-fly approach leads to 

efficient code (~320 source lines in Fortran), minimal 

storage requirements, and fast and consistent 

execution times. 

HYPAS consists of two phases as illustrated in 

Figure 10: 1) Capture Phase: Establishes pseudoequilibrium glide conditions; 2) Exit Phase: Exit 

conditions are predicted analytically assuming a 

constant altitude rate. The lift vector is adjusted to 

null the error between predicted and target apoapsis, 

and bank reversals are used to keep inclination errors 

within the desired limits. Results show excellent 

performance and an ability to capture ~98% of the 

theoretical corridor. 

PERFORMANCE/SIMULATION 

A Titan aerocapture simulation and Monte Carlo 

have been developed. All uncertainties critical to 

aerocapture robustness are included in the parameters 

varied in the Monte Carlo simulations. These 

parameters are the initial states with corresponding 

delivery errors at atmospheric interface; the 

atmosphere variability and uncertainty; including 

mean range and perturbations; aerodynamic 

uncertainties in CA, CN, αtrim, and cg uncertainties. 

During a Monte Carlo analysis, Monte Carlo 

parameters are each randomly varied over a specified 

range and distribution defined by the 

subsystem/model engineer. Two thousand trajectory 

simulations are completed in one Monte Carlo 

analysis. Results from these simulations provide 

aerocapture performance statistics to determine 

robustness, margin, risk; guidance development, 

stress case identification (control algorithm 

development – future); statistical distributions of 

critical parameters, design trajectories, for subsystem 

design. 

Monte Carlo analyses are completed for six Titan 

aerocapture mission scenarios. The first three are 

completed for the inertial arrival velocity of 6.5 

km/sec with no updates assumed from 

Figure 10. Aerocapture guidance phases. Ref 6. 

Figure 9. Effect of aerodynamic uncertainties on 

L/D. Ref 7.
7

Cassini/Huygens. This analysis is labeled “6.5 

km/sec” on the chart in Figure 11. As shown at the 

top of the figure, for this Monte Carlo 99.95% of the 

cases successfully aerocaptured at Titan; one of 2000 

cases failed. The first set of bars in Figure 11 show 

the corridor width required to accommodate aero 

uncertainty, atmosphere uncertainty, and delivery 

error, as well as the root sum square used to estimate 

total corridor width required for the 6.5 km/sec case. 

A comparison of the corridor width required and the 

theoretical corridor minus the loss in corridor due to 

guidance (only 2% loss), indicates that the 6.5 km/sec 

case with no update from Cassini/Huygens will be 

successful 3-sigma, but that there is not significant 

margin over and above a 3-sigma success. 

Adding the expected improvement in ephemeris from 

Cassini/Huygens in the next Monte Carlo analysis, 

labeled “6.5km/sec EOC ephemeris” shows the 

reduction in corridor required to accommodate 

delivery errors, and a corresponding reduction in the 

RSS 3-sigma corridor width required. The result is a 

significant increase in margin over and above a 3-

sigma success. These results are confirmed by the 

Monte Carlo analyses, where 100%, or 2000 of 2000, 

cases successfully aerocaptured. Further 

improvement results from the expected reduction in 

atmospheric uncertainty resulting from the 

Cassini/Huygens mission, labeled “6.5km/sec EOC 

ephemeris, atmos.” 

The next set of Monte Carlo’s completed is for 

10km/sec inertial entry velocity. The same three 

scenarios are completed for the 10km/sec entries as 

for the 6.5 km/sec entries. For each scenario, 100% 

of the Monte Carlo cases are successfully captured. 

Note that with increased velocity, the theoretical 

corridor increases faster than the net change in 

corridor width required due to aero, delivery and 

atmosphere errors and dispersions. Thus aerocapture 

performance robustness increases with velocity. The 

simulation, Monte Carlo analysis and results are 

discussed further in reference 7. 

AEROHEATING ENVIRONMENTS 

Conservative aeroheating design trajectories were 

selected for the reference concept. Design trajectories 

were based on the maximum atmospheric 

concentration of CH4, 5% by volume. The 

atmospheric CH4 concentration estimate is expected 

to either remain or be decreased with 

Cassini/Huygens data. 

A conservative ballistic coefficient, 90 kg/m2

, was 

assumed for the design trajectories. The resulting 

reference concept ballistic coefficient is 69kg/m2

, 

with a possible range of 56 – 84 kg/m2 for variation 

in vehicle diameter, aeroheating rates/loads, TPS 

selection and sizing (shown later in the paper). In 

addition, initial analyses show a decrease in 

aeroheating with decreasing ballistic coefficient, 

making 90 kg/m2

 conservative for the design 

trajectories. 

Lift-up trajectories were utilized to define the 

maximum heat rates; lift-down trajectories were used 

to define the maximum heat loads. These selections 

result in the most conservative trajectories for 

aeroheating environment definition. 

Based on the design trajectories and vehicle 

configuration, the peak laminar convective 

aeroheating rates are ~46 W/cm2 for 6.5 km/sec lift 

Figure 11. Titan aerocapture robustness. 
8

up, minimum atmosphere, alpha = 16 deg, stagnation 

point. Transition to turbulence is likely prior to peak 

convective heating on heatshield lee side based on an 

Retheta = 200 transition criteria and CFD results for 

the 6.5 km/sec lift up, minimum atmosphere, 16 deg 

angle-of-attack trajectory. This will likely increase the 

maximum convective heating on the heatshield beyond 

the above 46 W/cm2

.

9

For the maximum Titan CH4 concentrations, 5% by 

volume, all aeroheating predictions from the study 

have shown that the radiative aeroheating from CN is 

greater than the convective aeroheating. For the 

windside of the heatshield at alpha = 16 deg, the 

maximum radiative aeroheating is predicted to occur 

at the stagnation point. Radiative heating rates range 

from the “low end” ~93 W/cm2

 to the “conservative” 

~280 W/cm2

 for 6.5 km/sec lift up, minimum 

atmosphere, alpha = 16 deg, stagnation point. These 

results, completed during the study, were known to 

be preliminary. One of the recommendations from 

the study was to develop improvements in the 

radiative aeroheating environments analysis methods. 

This work is currently on going and discussed further 

in References 8 and 9. 

Based on the preliminary estimates, a range of 

radiative and convective heat rates and loads are 

estimated to provide a sensitivity of TPS selection 

and sizing to potential aeroheating environments for 

Titan aerocapture as shown in Figure 12. 

Figure 12. Aeroheating environments. 

TPS 

Candidate TPS materials are identified, and several 

sized, 10,11,12,13 for the range of Titan aeroheating 

environments expected. The ability of low density 

TPS materials (ex. SRAMs) to absorb radiation is 

currently unknown and requires testing. TUFROC, 

PhenCarb20, C-C Genesis-type are expected to 

absorb radiation. However these materials must also 

be tested in the radiative environments. In the TPS 

sizing analyses shown in Figure 13,10 it was assumed 

that the materials could absorb the radiation. 

Low-density silicone-based ablators are predicted to 

experience significant recession at the highest heating 

conditions. However they provide the lowest mass at 

lowest heating conditions. TUFROC and PhenCarb20 

exhibit less sensitivity to heating variations than low 

density silicone-based ablators, but with increased 

mass. PhenCarb 20 exhibits small surface recession 

in the inert Titan atmosphere. TUFROC is nonablating. C-C foam Genesis-type concept is also nonablating. 

For the Reference Concept, TUFROC (with the low 

end heat load sizing + 30%) was utilized in the MEL 

for two reasons. The TUFROC is expected to absorb 

radiation, and TUFROC TPS results in a conservative 

mass estimate compared to other candidate materials. 

TPS testing is needed to determine the ability of TPS 

materials to absorb CN radiation and to provide a 

family of TPS materials to accommodate a range of 

potential aeroheating environments at reduced mass. 

Figure 13. TPS thickness and mass vs. TPS type, 

aeroheating load. Ref 10.
9

AEROSHELL STRUCTURE 

The aeroshell structure was designed to be a current 

technology concept.14 Figure 1414 illustrates the 

structural components of the aeroshell. The loads are 

launch dominant with launch loads of 7 G axial, 3 G 

lat; and entry loads of 4 G axial with 3146 Pa on the 

heatshield for the 6.5 km/s entry. TPS masses used in 

the structure design are 1.181 g/cm2

 for the heatshield 

TUFROC TPS, and .187 g/cm2 for the backshell SLA 

TPS. 

The aeroshell was sized using 

HyperSizer/NASTRAN with the following results. 

The heatshield is 25.4 mm thick Hexcell 5052 Alloy 

Hexagonal Al Honeycomb core, with Graphite 

Polyimide facesheets. The backshell is 12.7 mm thick 

Hexcell 5052 Alloy Hexagonal Al Honeycomb core, 

with Graphite Polyimide facesheets. 

The aeroshell mass is summarized in Figure 15.14 The 

total aeroshell current best estimate mass (structure + 

TPS + non structural) is 327.80 kg. 

MASS PROPERTIES, PACKAGING 

Figure 16 shows the stack and orbiter packaging 

design.4

 Note that the aeroshell size and packaging 

efficiency are governed by the 2.4 m diameter high 

gain antennae (HGA) packaging. 

Figure 17 shows a system level mass summary for 

the Delta 4450, SEP, Earth Gravity Assist (EGA), 

aerocapture concept. The concept has 30% system 

level margin, and greater than 10% system reserve. A 

VGA option with a Delta 4450, SEP, VGA, 

aerocapture has 6% system reserve. The aerocapture 

mass fraction is 41.5% of orbiter entry mass. These 

results are not possible without this level of detail in 

packaging, s/c design, structure, and TPS. 

Backshell

Pallet 

Ring

Heatshield

Separation

Mechanism

(6 locations)

Backshell

Pallet 

Ring

Heatshield

Separation

Mechanism

(6 locations)

Figure 14. Aeroshell structural components. Ref 14. 

Component Area 

(m2

) 

Structure 

Mass 

(kg) 

TPS 

Mass 

(kg) 

Non 

Structural 

Mass (kg) 

Heatshield 12.58 41.58 148.62* 0 

Backshell 15.01 43.27 28.69** 2.38 

Pallet Ring 1.20 42.47 0 1.20 

Sep Ring 1.79 11.35 0 .89 

Sep Ring 

Attachments 0.45 2.85 0 4.50 

Totals 141.52 177.31 8.97 

Figure 15. Aeroshell structure mass. Ref 14. 

SEP Prop

Module

Solar

Arrays

Orbiter

Lander

3.75 m diameter

Aeroshell

2.4 m diameter HGA

Figure 16. Refererence mission packaging. Ref 4. 

Mass (kg)

Component

Current Best 

Estimate

% 

Contingency Growth

System 

Allocation

Lander 280.2 29.8% 363.8 400.0

Orbiter/Lander Interface 47.5 30.0% 61.8 61.8

Orbiter 883.6 24.2% 1097.7 1200.0

Prop Mod/Orbiter Interface 47.3 30.0% 61.4 61.4

SEP Prop Module 1084.0 21.4% 1316.5 1450.0

Launch/Prop Mod Interface 60.0 30.0% 78.0 78.0

Stack Total 2402.6 24.0% 2979.2 3251.2

3423

29.8%( LV Cap - CBE ) / LV Cap

13.0%( LV Cap - Growth ) / LV Cap

Subsystem Rack-up

System Reserve

System Level Mass Margin

Launch Vehicle Capability

Figure 17. System level mass properties for Reference 

Titan Explorer aerocapture mission. Ref 4. 
10

MASS PROPERTY SENSITIVITIES 

Several mass property sensitivity analyses were 

completed to determine aerocapture mass fraction, 

aeroshell mass, system mass sensitivity to 

aeroheating environment assumptions, TPS 

candidates and aeroshell size.15 An assessment of 

aerocapture system mass to an all propulsive mass 

system was also completed to determine the mass 

savings for an aerocapture system.15, 3 

The aerocapture system mass fraction as a function of 

aeroheating environment and TPS concept is shown 

in Figure 18.15 For example, SRAM-20 is sized for 

three levels of aeroheating environments. Results are 

labeled SRAM-20, SRAM-20 +1, SRAM-20 +2, for 

the low, medium and conservative aeroheating levels 

respectively. The potential savings in mass for 

alternate TPS concepts, compared to the baseline, is 

evident. In addition, the effect of aeroheating 

environment on TPS mass is also shown. The 

ballistic coefficient range is 56 – 72 kg/m2

 over this 

range of aeroheating environments and TPS materials 

for the 3.75 m diameter aeroshell. 

Based on the orbiter packaging, a reduction in the 

aeroshell diameter may be possible. The minimum 

diameter would be approximately 3 m. Further 

reduction would preclude packaging of the spacecraft 

with the 2.4 m HGA. A maximum of 20% reduction 

in aerocapture system mass fraction results for a 

minimum 3 m diameter aeroshell. The ballistic 

coefficient range for aeroshell size range of 3 – 3.75 

m diameter and range of TPS is 56 – 84 kg/m2

, all 

lower than the design trajectory ballistic coefficient 

of 90 kg/m2

. 

Titan Aerocapture System Mass Fraction vs Forebody TPS Material: 3.75 

m Diameter, 6.5 km/s Entry, SLA B/S TPS

27%

29%

31%

33%

35%

37%

39%

41%

43%

45%

SRAM-14SRAM-17

SLA-561V

SRAM-20

SRAM-20 +1

SRAM-20 +2

PC-20

PC-20 +1

PC-20 +2

TUFROC

TUFROC +1

TUFROC +2

C-C

C-C +1

C-C +2

Forebody TPS Material

Aerocapture System Mass Fraction

50% Margin

30% Margin

Nominal

Thickness

Genesis Style + 15%

Reference MEL with contingency =38.8% of Orbiter Launch Wet Mass

 = 41.5% of Aerocapture Mass

Based on 426.14 kg aerocapture system mass.

* Baseline TPS

*

*

Figure 18 Aerocapture system mass fractions vs. forebody TPS material and aeroheating environments. 

Figure 19 Aerocapture vs. All-Propulsive. *Includes 2 year moon tour used to reduce the propellant 

requirements for all propulsive capture. 
11

Figure 19 shows a comparison of the Titan Explorer 

mission with aerocapture compared to an allpropulsive option. For the Titan Explorer payload 

mass, the all-propulsive option was not feasible with 

a Delta 4450, while aerocapture enables a feasible 

mission on the 4450. On a Delta IV Heavy an allpropulsive option is marginal, but potentially 

feasible. However trip times are 12 years for an all 

propulsive option compared to 6 years for an 

aerocapture mission. In addition the aerocapture 

mission is capable of delivering 2.4 times more 

payload to Titan compared to the all-propulsive 

mission for the same launch vehicle. Dependent on 

mission requirements including required payload 

mass and trip time, for example, aerocapture is 

enabling to strongly enhancing compared to an allpropulsive mission. 

SUMMARY 

Aerocapture is feasible, and the performance is 

adequate, for the Titan mission with the high heritage 

low L/D configuration and no improvements from the 

Cassini/Huygens mission. 99.95% success results for 

a 6.5 km/sec entry with the low L/D configuration, 

but with low margin above 3-σ. 100% success results 

for a 10 km/sec entry with the low L/D configuration, 

with margin above 3-σ success. Cassini/Huygens 

results in an increase in margin and robustness 

significantly above 3-σ, for 6.5 km/sec and 10 km/sec 

entry aerocapture at Titan with the high heritage low 

L/D configuration. 

Aerocapture can deliver 2.4 times more mass to Titan 

than an all-propulsive system for the same launch 

vehicle. 

TECHNOLOGIES 

Technologies identified in the study as needing 

development were grouped into three categories; 

enabling technologies, strongly enhancing 

technologies and enhancing technologies. As noted 

earlier, aerocapture is enabling to strongly enhancing 

for the Titan missions, dependent on the mission 

requirements. However aerocapture is ready for an 

Earth or Mars flight experiment without additional 

technology development. An aerocapture flight 

experiment may be desirable to demonstrate 

aerocapture technology to reduce risk/cost for science 

mission acceptance. 

The strongly enhancing technologies identified 

include: 

• Aeroheating methods development and 

validation. Large uncertainties currently exist, 

improved prediction capability could result in 

reduced TPS mass, as well as support 

configuration trades and further mission design 

trades. 

• TPS material testing. TPS materials proposed 

and other TPS options exist today, but have not 

been tested against the expected radiative 

heating at Titan. 

• Atmosphere Modeling. Titan General 

Circulation Model output is needed to represent 

the “true” natural variability of the atmosphere. 

• The enhancing technologies identified include 

• Aeroshell lightweight structures for reduced 

aerocapture mass. 

• Guidance – Existing guidance algorithms have 

been demonstrated to provide acceptable 

performance, improvements could provide 

increased robustness. 

• Simulation – Huygens trajectory reconstruction, 

statistics and modeling upgrades. 

• Mass properties/structures tool for systems 

analysis capability improvement, concept 

trades. 

• Deployable high gain antennae for increased 

data return. 

• The following technologies provide significant 

benefit to the mission but are already in a 

funded development cycle for TRL 6 by 2006. 

• MMRTG (JPL sponsored AO in proposal 

phase, First flight Mars ’09) 

• SEP engine (Glenn Research Center engine 

development complete in ‘0#) 

• Second Generation AEC-Able UltraFlex Solar 

Arrays (175W/kg) 

• MRO optical navigation camera to be 

demonstrated in MRO. 

ADDENDUM 

Aerocapture Characteristics: Titan vs. Mars 

Since much of the higher fidelity aerocapture work 

has been done for Mars missions, Figure 20 provides 

a comparison of a representative Mars aerocapture 

mission to the Titan aerocapture reference mission 

described in this paper for reference. 
12

Destination/ Reference 

Parameter Titan Mars 

Entry Velocity (km/sec) 6.5* 5.7 

Nom. Entry Flight Path 

Angle (deg) -36 -14.2 

Apoapsis/Science Orbit 

(km) 1700 1400 

Atmosphere Composition 95%N2, 

5%CH4** 

95.3% 

CO2, 

2.7%N2 

Atmos Scale Height at 

Aerocapture Altitude ~40 10.5 

Atmospheric Interface 

Altitude (km) 1000 250 

Aerocapture Altitude (km) 200-400 40 

L/D .25 .25 

M/CDA (kg/m2

) 90*** 148 

Theoretical Corridor (deg) 3.5 ~1.4 

Time from Atmos Entry to 

Atmos Exit (min) 42 10 

Convective Stagnation 

Point Heat Rate (W/cm2

) 

46 (.91m 

nose rad) 

30 (1.9m 

nose 

radius) 

Radiative Aeroheating 

Rate (W/cm2

) 93-280 negligible 

Max g’s During 

Aerocapture (Earth g’s) 3.5 2.5 - 3 

Figure 20. Representative aerocaptures at Titan 

compared to Mars. *Titan aerocapture entry 

velocity 6.5 – 10km/sec, comparison given for 

6.5km/sec. **Max CH4 atmosphere. ***For 

design trajectory & comparison; range: 56–84 

kg/m2

 dependent on aeroheating, TPS, vehicle 

diameter 

REFERENCES 

1. “Titan Lander Conservative Science 01-06” 

results dated 4,5,8,15 June, 2001. 

2. Haw, R. “Approach Navigation for a Titan 

Aerocapture Orbiter,” AIAA-2003-4802. 

3. Noca, M., Bailey, R., “Titan Explorer Mission 

Trades from the Perspective of 

Aerocapture,”AIAA-2003-4801. 

4. Bailey, R., Hall, J., Spilker, T., “Titan 

Aerocapture Mission and Spacecraft Design 

Overview,” AIAA-2003-4800. 

5. Justus, J., Duvall, A., “Engineering-Level 

Model Atmospheres for Titan and Neptune”, 

AIAA-2003-4803. 

6. Masciarelli, J., “Guidance Algorithms for 

Aerocapture at Titan,” AIAA-2003-4804. 

7. Way, D., Powell, R., Edquist, K., “Aerocapture 

Simulation and Performance for the Titan 

Explorer Mission,” AIAA-2003-4951. 

8. Olejniczak, J., Prabhu, D., Wright, M., 

Takashima, N., Sutton, K., “An Analysis of the 

Radiative Heating Environment for Aerocapture 

at Titan,” AIAA-2003-4953. 

9. Takashima, N., Hollis, B., Olejniczak, J., 

Wright, M., Sutton, K., “Preliminary 

Aerothermodynamics of Titan Aerocapture 

Aeroshell,” AIAA-2003-4952. 

10. Laub, B., “Thermal Protection Concepts and 

Issues for Aerocapture at Titan,” AIAA-2003-

4954. 

11. Curry, D.M., Congdon, W.M., “2007 Titan 

Mission Analysis of ARA Ablators to Define 

Required Thickness and Weights,” Report 

presented to Titan Aerocapture Systems 

Analysis Team, Report Date 11July2002. 

12. Congdon, W.M., “TPS-Table-Titan,” e-mail to 

M.K.Lockwood, 23Aug2002. 

13. Dec, J.A., “Titan Orbiter Heatshield Sizing 

Study,” e-mail to M.K.Lockwood, 27Sep2002. 

14. Hrinda, G., “Structural Design of the Titan 

Aerocapture Mission,” AIAA-2003-4955. 

15. Dyke, R. E., “Planetary Probe Mass Estimation 

Tool Development and Its Application to Titan 

Aerocapture,” AIAA-2003-4956. 

Acknowledgements: The author would like to 

acknowledge and thank the team members of the 

NASA Aerocapture Systems Analysis Team for their 

work and contributions to the Titan Aerocapture 

Systems Analysis Study and to this paper. Thank you 

to Carol Carroll, Aerocapture Systems Analysis 

Study Review Panel Chair, and the review panel 

members for review of this work and helpful 

comments and recommendations. Thank you to Anne 

Costa for preparing this paper for publication. 
13 

TITAN AEROCAPTURE MISSION AND SPACECRAFT DESIGN OVERVIEW 

R. W. Bailey, J. L. Hall, T. R. Spilker 

Jet Propulsion Laboratory 

California Institute of Technology 

ABSTRACT 

A detailed Titan aerocapture systems analysis and 

spacecraft design study was performed as part of 

NASA’s In-Space Propulsion Program. The primary 

objective was to engineer a point design based on 

blunt body aeroshell technology and quantitatively 

assess feasibility and performance. This paper 

provides an overview of the mission and spacecraft 

design resulting from that study and references other 

papers that provide further details on critical 

subsystems. It also reviews the science requirements 

underlying the selected mission concept of an 

aerocaptured orbiter and a separate entry vehicle that 

delivers an aerobot into the Titan atmosphere. 

Including aeroshells and 30% contingencies, the 

estimated mass of the orbiter is ~1100 kg and that of 

the entry vehicle ~360 kg. Solar electric propulsion 

(SEP) and an Earth gravity assist is used to get the 

tandem vehicle to Titan in 6.5 years, with orbiter – 

entry vehicle separation occurring one month prior to 

arrival. The SEP module, orbiter and entry vehicle 

are vertically stacked on a medium class launch 

vehicle and connected with a truss structure. Power 

profiles based on a strawman instrument suite and 

telecom strategy are accommodated with a pair of 

120 W (electric) radioisotope thermoelectric 

generators. Details on the configuration layout, mass 

and power breakdowns, key design trades and 

outstanding design issues are also included. 

TABLE OF CONTENTS 

1. Introduction 

2. Science 

3. Mission Overview 

4. Mission System Description 

5. Atmospheric Probe Design 

6. SEP Module Design 

7. Orbiter Design 

8. New Technology Development 

9. Recommended Additional Analysis 

10. Conclusions 

11. Acknowledgements 

12. Bibliography

INTRODUCTION 

As part of NASA’s In-Space Propulsion Program, 

aerocapture is being investigated as a means for 

interplanetary orbit insertion. A systems analysis and 

spacecraft point design study was performed in the 

Fiscal Year 2002 time frame based on a reference 

mission to Saturn’s moon Titan. The purpose of this 

study was to quantify the feasibility and performance 

of an aerocapture system to insert a spacecraft into a 

science orbit about Titan. This paper provides an 

overview of the mission and spacecraft design 

resulting from that study and references other papers 

presented at this conference that provide further 

details on mission design, navigation, critical 

subsystems and the aerothermal environment for 

aerocapture at Titan. 

The overall mission concept includes the delivery of 

a long duration atmospheric probe to Titan’s 

atmosphere and the use of a Solar Electric Propulsion 

(SEP) stage for the Earth to Saturn transit. The 

mission concept is shown to be feasible at the level of 

detail applied for this study. Many different technical 

areas and trades consistent with continued Phase A/B 

efforts are defined at the end of this paper. 

SCIENCE 

Objectives & Measurements 

For this study, the primary science objectives were 

taken from Chyba et al.1

 for a post-Cassini / Huygens 

Titan mission, listed in priority order: 

1. Distribution and composition of organics 

2. Organic chemical processes, their chemical 

context and energy sources 

3. Prebiological or protobiological chemistry 

4. Geological and geophysical processes and 

evolution 

5. Atmospheric dynamics and meteorology 

6. Seasonal variations and interactions of the 

atmosphere and surface (not addressed in a 

mission of short lifetime) 

These objectives will likely be revisited when results 

are available from the Cassini / Huygens mission. 

Tamppari et al.

2

 involved the Titan science 

community in a workshop that prioritized 
14 

measurement objectives for such a mission. The 

highest priorities were determined to be: 

1. Global surface morphology 

2. Global gross surface composition and chemistry 

3. Atmospheric composition and its spatial and 

temporal variability 

4. Atmospheric structure and its spatial and 

temporal variability: vertical profiles of density, 

pressure, and temperature 

5. Atmospheric dynamics (winds) and meteorology 

An independent external review performed after the 

completion of this study judged these objectives to be 

appropriate. 

Science Instruments 

Table 1 presents the instrument suite selected and the 

flowdown from science and measurement objectives 

to the instruments. Although these instruments may 

be realistic, for the purpose of this study they serve as 

mass, power, and data volume placeholders for A 

Titan science payload. 

Multi-Spectral Imager 

The multi-spectral imager uses spectral coverage in 

several atmospheric opacity “windows” between 1 

and 5 microns to determine surface and atmospheric 

morphology and chemistry as well as atmospheric 

dynamics and meteorology. This instrument will fill 

in any coverage gaps remaining after Cassini / 

Huygens. 

Synthetic Aperture Radar (SAR) 

The SAR uses the Orbiter X-Band telecom system 

with the HGA pointed off-nadir. It makes 

complementary measurements of surface morphology 

and meteorology, through clouds that would obscure 

the imaging instruments’ view, and can detect the 

bottoms of shallow hydrocarbon lakes. Like the 

imager, it will fill in any coverage gaps remaining 

after Cassini / Huygens. 

Microwave Spectrometer 

The microwave spectrometer, capable of either nadiror limb-pointed modes, makes global, low (spatial) 

resolution measurements of atmospheric structure, 

dynamics, and meteorology via detailed spectroscopy 

of emission lines from a few key chemical species. 

This also yields precise vertical abundance profiles of 

those species. 

Ultrastable Oscillator (USO) 

Adding a USO to the Orbiter X-Band telecom system 

enables atmospheric radio occultation science. 

Radiometrics obtained when the signal path to Earth 

passes through Titan’s atmosphere allows accurate 

(1-2%), high-resolution vertical profiling of 

temperatures and densities at many sites, yielding 

atmospheric structure and dynamics as well as 

ionospheric structure 

Atmospheric Probe Science 

The Atmospheric Probe (AP) was allocated 5.3 Gbits 

of total data return; or the capability of the UHF relay 

link over a one year period. The AP to Orbiter link 

provides 64 kbps for 30 minutes every 8 days. 

Although 5.3 Gbits is adequate for general 

atmospheric and meteorological data (~14 

Mbits/day), this volume is likely inadequate for any 

type of context imaging – this is generally an issue 

for the AP and not addressed in this study which 

focuses on the aerocapture technology aspects of the 

mission. 

Table 1. Science Instruments 

Instrument 

CBE 

Mass 

(kg) 

CBE 

Power 

(W) 

Spatial 

Res 

(meters) Point 

FOV 

(deg) 

Measure 

Objective 

Science 

Objective 

Total 

Data 

Return 

(Tbits) 

Multi-spectral Imager 12 14 ~30 Nadir 1.0 1,2,3,5 1,2,3,5 9 

Synthetic Aperture Radar 10 30 ~200 OffNadir 1.15 1,2 4 1.9 

Microwave Spectrometer 10 50 N/A Nadir & 

Limb 

6.5 3,4,5 2,3,5 1.0E-4 

Ultra Stable Oscillator 0.8 3 N/A N/A N/A 4,5 5 N/A 
15 

MISSION OVERVIEW 

The study was based on a Titan Explorer concept with 

an Orbiter and an Atmospheric Probe (AP). Certain 

aspects of the mission were assumed as ground rules 

from previous studies performed internally at the Jet 

Propulsion Laboratory’s (JPL) Team-X12. Other aspects 

of the mission were open to system trades and/or 

inherited from other outer planet mission studies 

performed internally at JPL. 

Ground Rules 

Several ground rules and assumptions were set to bound 

the study. These items were not subject to any system 

trades analysis. 

• The mission shall deliver an AP into the Titan 

atmosphere, and a spacecraft into Titan orbit. 

• The total mission lifetime shall be no longer than 

10 years. 

• The Technology Readiness Level 6 cutoff date 

shall be no later than Dec 2006. 

• The AP will be a “black box” with a 400 kg launch 

mass allocation. 

• The AP operational lifetime will be 1 year. 

• The Orbiter shall perform an aerocapture for Titan 

orbit insertion. 

• The Orbiter shall provide global coverage 

opportunity for all the science instruments. 

• Science data return shall utilize no more than 8 

hours per day of a 70m ground station. 

Earth to Saturn Trajectory 

The Earth to Saturn trajectory, shown in Figure 1, 

provides a good combination of transit time, Titan entry 

velocity, launch mass, and SEP propellant/power mass. 

Many different trajectories were considered which 

included different launch vehicles, launch dates, transit 

times, SEP power levels, number of SEP ion engines, 

and planetary gravity assists. These trajectory options 

and their associated trades are discussed in detail by 

Noca, et al.

3

 The important aspects of the selected 

trajectory are as follows: 

 Launch Vehicle: Delta IV M (4450-14) 

 Launch C3: 8.6 km2

/sec2

 Launch Mass: 3423 kg (10% reserve) 

 Launch Date: Dec 24, 2010 

 Gravity Assist: Earth 

 SEP Burn Time: 30 months, accumulated 

 SEP Power: 24 kW (End Of Life) 

 SEP Propellant: 460 kg (no contingency) 

 Transit Time: 5.9 years 

 Titan Entry Velocity: 6.5 km/sec 

Figure 1. Earth to Saturn SEP trajectory 

Mission Timeline 

The mission timeline is listed below. For the “Time” 

column, ‘L’ = Launch, ‘A’ = Orbiter atmospheric 

interface, ‘y’ = years, ‘d’ = days, ‘h’ = hours, and ‘m’ = 

minutes. 

Time Event 

L+0 Launch, SEP burn start 

L+23m Earth flyby 

L+30m SEP burn out and jettison at ~2.5 AU 

L+5.7y (A-60d) Traj Correction Maneuver (TCM) 1 

A-31d Probe Release TCM (2) 

A-30d Probe Release 

A-29d Post Release TCM (3) 

A-7d TCM 4 

A-1d TCM 5 

A-6h TCM 6 (if needed) 

A-3h Probe entry 

A-1h Jettison non-aero external components 

A-30m Align for aerocapture interface 

A+20m Jettison aeroshell 

A+4h Periapsis raise (circularization) burn 

A+3y End of mission 

Once the Solar Electric Propulsion Module’s (SEPM) 

job is done at around 2.5 AU, it is jettisoned to 

eliminate mass and solar array perturbations for later 

TCMs. The Orbiter uses a combination of Doppler 

ranging, ∆DOR, and optical navigation4

 to setup the AP 

entry trajectory delivery at entry minus 30 days. 
16 

Figure 2 illustrates the final aerocapture trajectory. The 

Orbiter spins up providing the AP with attitude 

stabilization, and then separates the AP. The Orbiter despins and performs a separation maneuver designed to 

put 3 hours of separation between the AP and Orbiter 

atmospheric entries. This allows the Orbiter to receive 

AP critical event data during AP entry, descent, and 

initial checkout. The Orbiter relays this data to Earth 

before it enters Titan’s atmosphere. Approximately 1 

hour prior to Orbiter atmospheric entry, the Orbiter will 

eject all non entry system components (truss, radiators, 

antennas, etc), and then orient for entry. 

Figure 2. Aerocapture Trajectory 

The primary heat pulse of aerocapture lasts less than 10 

minutes, during which the Orbiter is actively 

controlling its bank angle with hydrazine thrusters. 

After atmospheric exit the aeroshell is jettisoned and 

the Orbiter prepares for the periapsis raise maneuver to 

insert the Orbiter into a 1700 km circular orbit. 

MISSION SYSTEM DESCRIPTION 

The Orbiter Flight System is the primary focus of this 

study and the Ground Data and Mission Operations 

Systems were not addressed. The SEPM is largely 

inherited from previous study and the AP is treated as 

a black box. The Launch and SEP configurations are 

shown in Figures 3 and 4, respectively. The launch 

system mass summary is shown in Table 2. The post 

SEP cruise configuration is shown in Figure 5. 

Figure 3. Launch Configuration 

Figure 4. SEP Burn Configuration 

Figure 5. Post SEP Cruise Configuration 
17 

Table 2. Launch Mass Summary (kg) 

CBE = Current Best Estimate 

GC = Growth Contingency = ( GE - CBE ) / CBE 

GE = Growth Estimate 

A = Allocation from system 

Launch Dry Mass Margin = same as Launch Wet 

Margin with total propellant mass subtracted from all 

estimates. 

Key Mission System Trades 

Several trades associated with the overall Flight System 

are worth mentioning. These trades do not represent the 

a complete trade space for the Titan mission only those 

trades which drove the configuration of the flight 

system to allow convergence of a mission concept and 

determine aerocapture system feasibility. 

Launch Stack Configuration 

The orientation and placement of the AP and the 

Orbiter on top of the Propulsion Module (PM) drove 

the structural mass of the adapters as well as the 

primary structure mass for the PM and the orbiter. In 

general, 4 configurations were analyzed: the Orbiter 

and AP in nose up and nose down configurations with 

the Orbiter below and above the AP. The final 

configuration selected was Orbiter nose down (with 

respect to the launch vehicle) below the AP oriented 

nose up. 

For all configurations of the Orbiter above the AP, the 

PM to Orbiter adapter became complex and massive. 

Additionally the large Orbiter mass suspended high on 

the launch stack resulted in much higher SEPM 

structure mass to accommodate the lateral launch loads 

and frequencies. 

The Orbiter / AP orientation was selected as tail to tail 

for two reasons. First, there are potentially three 

separation planes between the Orbiter and the AP: 

Orbiter/AP, Orbiter/Truss, and AP/Truss. Each 

separation plane poses scarring risks to the TPS of the 

Orbiter and AP. Second, the structural interface of the 

truss with each vehicle is a risk to the TPS burn through 

(interface results in localized thermal anomalies). 

Routing the structural interface through the aft body 

lowers these risks. 

The resulting stack configuration routes the primary 

PM to Orbiter adapter structure through the Orbiter fore 

body TPS. An engineering solution to the localized 

thermal anomalies at the interface with the TPS is 

considered solvable, but at the same time, it is 

considered highly desirable to find an alternate 

configuration which avoids perforation of all fore body 

TPS. 

Cruise Propulsion System 

A SEPM was selected over a chemical stage for the 

Earth to Saturn trajectory because of the SEP 

trajectory’s superior overall performance in terms of 

delivered mass, flight time less than 6 years, and 

atmospheric entry velocity of around 6.5 km/sec. A 

detailed discussion of the chemical versus SEP trade is 

addressed by Noca, et al.

3

AP / Orbiter Delivery 

The delivery of the AP and the Orbiter to their 

respective entry trajectories could be performed by the 

Orbiter or by the SEPM. Since the aerocapture phase 

required the Orbiter to have all the subsystems required 

to perform AP and Orbiter delivery, the Orbiter was 

selected to perform these entry trajectory deliveries. 

This allowed deletion of the ACS, C&DH, and telecom 

subsystems from the SEPM and the separation of the 

SEPM soon after its burn out. 

Probe Entry and Descent Data Relay 

The AP critical event relay during entry and descent is a 

multi dimensional trade involving delivery errors 

(Orbiter and AP), telecom (AP and Earth), and Orbiter 

atmospheric entry risk (late separation of non entry 

system components). The selected strategy may not be 

the best solution, but is adequate to show feasibility for 

this study. The aspects of the Probe to Orbiter relay link 

are discussed in more detail in a later section. 

Critical events relay using the SEPM on a flyby 

trajectory was ruled out because this would require the 

SEPM to be an independent spacecraft with 

unnecessary functional duplication with the Orbiter. 

Element CBE GC GE A

Atmospheric Probe 280.2 29.8% 363.8 400.0

Orbiter/AP Interface 47.5 30.0% 61.8 61.8

Orbiter Dry Mass 743.0 28.4% 954.0

Orbiter Prop Mass 140.6 2.2% 143.7

SEP/Orbiter Interface 47.3 30.0% 61.4 61.4

SEP Dry Mass 623.9 29.9% 810.4

SEP Prop Mass 460.1 10.0% 506.1

Launch/SEP Interface 60.0 30.0% 78.0 78.0

Dry Mass (DM) Totals 1801.9 29.3% 2329.4 2601.4

Prop Mass (PM) Totals 600.7 8.2% 649.8 649.8

Stack Total 2402.6 24.0% 2979.2 3251.2

Launch Vehicle Capability 3423 LVC

Launch Wet Mass Margin 29.8% ( LVC - CBE ) / LVC

System Reserve 13.0% ( LVC - Growth ) / LVC

Launch Dry Mass Margin 35.0%

1200.0

1450.0
18 

Titan Orbit Altitude 

The initial desired science orbit altitude was specified at 

1400 km. Orbit maintenance analysis performed by 

LaRC showed that for ballistic coefficients similar to 

the Orbiter design, as much as 100 m/s would be 

required for a three year mission – resulting in more 

propellant mass than the Orbiter could carry in our 

reference mission. The same LaRC analysis showed 

less than 2 m/s if the altitude was raised to 1700km. 

The only impact from a higher orbital altitude was to 

the science instruments. Since none of the specific 

instruments exist at this time, it was determined that the 

instrument impact was acceptable. Actual atmospheric 

density results from the Hyugens probe may provide an 

opportunity to lower this altitude if necessary. 

ATMOSPHERIC PROBE DESIGN 

As stated earlier, the AP design is considered to be a 

black box and out of the scope of this study. There are 

internal JPL studies performed by Team-X12 indicating 

that 400kg is an adequate allocation for a Titan AP. 

SEP MODULE DESIGN 

A SEPM was selected over a chemical PM for the 

overall combination of shorter flight time, lower entry 

velocity, and lower PM mass; this trade is discussed in 

more detail by Noca3

. An existing JPL SEPM design 

was modified for the Titan mission. The primary 

modifications were deleting the avionics in favor of 

using the Orbiter’s avionics and increasing solar array 

structure and power capability. The SEPM structural 

mass was analyzed to assure proper launch load and 

frequency capability for the entire launch stack. The 

SEPM mass summary is shown in Table 3. 

Table 3. SEP PM Mass Summary (kg) 

Flt 

Unit CBE GC GE

1084.0 21% 1316.5

460.10 10% 506.11

56 623.93 30% 810.40

ACS, C&DH, Telecom 3 5.85 30% 7.61

Power 7 148.31 30% 192.80

Structure 14 281.39 30% 365.81

Propulsion 20 142.28 30% 184.97

Thermal 12 46.10 28% 59.21

SEP Dry Mass

Xenon Propellant

SEP Wet Mass (kg)

Element

Most of the Attitude Control System (ACS), Command 

& Data Handling (C&DH), and Telecom functionality 

was moved to the Orbiter or to the Orbiter / Probe 

structural interface. This eliminates unnecessary 

component duplication through the entire system and 

allows the SEPM to be jettisoned after its job is 

complete. The majority of the Power system mass is the 

25.6 m2

, 178 kg (growth), of solar arrays. The “Ftl 

Unit” column of Table 3 specifies the number of line 

items in the detailed mass list for the respective 

subsystem. 

ORBITER DESIGN 

The Orbiter design is shown in Figures 6, 7 and 8. 

Figure 6 shows the Orbiter in the Post AP release 

configuration. This configuration shows the critical 

components required by the Orbiter through post 

launch, SEP cruise, and AP entry and descent. These 

elements include Orbiter electronics and MMRTG 

radiators, X-Band MGA, AP UHF Relay antenna, and 

optical navigation cameras. Figure 7 shows the 

aerocapture configuration and Figure 8 specifies the 

primary Orbiter components. 

Figure 6. Post AP Release configuration 

Figure 7. Aerocapture Configuration 
19 

Figure 8. Titan Orbit Configuration 

The estimated power required for the various mission 

phases is summarized in Table 4. Heater power in all 

phases is minimal because of an assumption that the 

MMRTG excess heat, ~3700W, can be distributed 

across the spacecraft well enough to not require the 

heater power typical for deep space missions. The 

modes listed in Table 5 are not all the modes identified 

in the study, just the ones that stress the system. The 

available power listed is the power output of two 

MMRTGs after 1.5% output degradation per year. It is 

assumed that once the telecom and instrument 

components have been turned on, that they are never 

turned completely off, but rather are placed in a low 

power standby mode when not in use. The 25% margin 

shown in the table is typically considered not viable for 

a pre-project, but specific opportunities to improve this 

are noted in the power subsystem section. 

Table 4. Orbiter Average Power (W) 

Mission Phases

Cruise

Aero 

Capture

Lander 

Relay

Orbit 

Science

Earth 

Comm

Margin 38.1% 46.6% 32.4% 26.9% 25.0%

Available 226 226 222 214 214

Totals 140.00 120.60 150.00 156.50 160.50

Instrument 1.5 1.5 18.0 70.5 18.0

ACS 40.1 44.2 40.1 40.1 40.1

C&DS 14.9 14.9 14.9 14.9 14.9

Power 3 33 33

Propulsion 0 27 0 0 0

Telecom 76.5 30.0 66.0 20.0 76.5

Thermal 4.0 0.0 8.0 8.0 8.0

Table 5 presents the Orbiter mass summary with 

subtotals for Titan orbit, entry, and launch mass. All 

components are block redundant except for structure, 

propulsion, antennas, MMRTGs, thermal radiators, and 

science instruments. Generally, the Orbiter design was 

driven towards single fault tolerance without mission 

loss. Generally, a growth contingency (GC) of less than 

30% in Table 5 indicates where components of high 

heritage are utilized in the system design. Instruments, 

structure, power, and thermal are the primary new 

development subsystems. 

Table 5. Orbiter Mass List Summary (kg) 

Flt 

Unit CBE GC GE

883.6 24% 1097.7

140.6 10% 143.7

293 743.0 28% 954.0

Cruise Support Mass 65.5 9% 71.6

41.4 0% 41.4

49 24.1 25% 30.2

Structure & Misc 7 8.2 16% 9.5

Thermal 42 15.9 30% 20.7

818.1 25% 1026.1

416.7 24% 515.0

88.9 0% 88.9

Bank Angle Control 12.6 0% 12.6

Circularization dV 76.3 0% 76.3

8 327.8 30% 426.1

TPS 2 177.3 30% 230.5

Structure 6 150.5 30% 195.6

401.4 27% 511.1

10.3 30% 13.4

236 391.1 27% 497.7

Instruments 4 32.8 30% 42.6

ACS 15 20.2 10% 22.2

C&DH 16 15.3 26% 19.2

Power 5 80.1 30% 104.1

Telecom 13 39.1 24% 48.4

Structure 7 136.7 30% 177.7

Propulsion 73 39.5 21% 47.8

Thermal 103 27.4 30% 35.7

Dry Mass

Propellant

Dry Mass

Titan Orbit Wet Mass

Propellant

Propellant

Dry Mass

Atmos Entry Wet Mass

Aerocapture System

Total Dry Mass

Total Propellant

Launch Wet Mass

Orbiter Element

Key Flight System Trades 

Reaction Wheels vs. Thrusters 

The approach navigation and the science teams would 

prefer reaction wheels for attitude control rather than 

RCS jets. Unfortunately, the 2 MMRTG design could 

not supply enough power to operate 3 reaction wheels 

and everything else needed for science data gathering. 

Minimum Impulse Thrusters (MIT), TRL 6 in 2004, 

were selected as a compromise providing the necessary 

stability for science pointing, but degraded performance 
20 

for approach navigation (small forces integration). The 

navigation analysis performed by Haw4 considers a 

spacecraft with reaction wheels; this inconsistency was 

not resolved before the end of the study. 

Rigid vs. Deployable X-Band HGA 

The selected 2.4m Fixed HGA selected provides a 2.3 

Tbit data return capability to a 70m station assuming 

one 8 hour contact every day. This falls well short of 

the ~11 Tbits generated by the science instruments. A 6 

meter deployable HGA was investigated, but it was not 

able to be incorporated into the design before the end of 

the study. Follow-on work for a Neptune aerocapture 

mission has since determine a deployable HGA is a 

feasible concept for the Titan mission timeframe. A 6 

meter antenna would provide a 500 kbps return link 

capable of returning 15 Tbits and provide an 

opportunity to reduce the aeroshell diameter and change 

the backshell design to a single angle. Such an antenna 

would increase pointing knowledge and control 

requirements, but these requirements are to be within 

the capability of the MIT RCS based ACS system. The 

larger antenna would also affect the SAR instrument 

design because of the narrower beam width – higher 

resolution, but possible less than global coverage. 

Subsystem Descriptions 

Orbiter subsystems will be discussed in order of overall 

system impact. In general, subsystems discussed first 

drive the system design more than those discussed last. 

Aerocapture System 

The aerocapture system is defined as the TPS, the 

underlying aeroshell structure, the propellant required 

for attitude control during aerocapture, and the 

propellant required for the orbit circularization burn at 

the apoapsis of the aerocapture exit orbit. 

The aerocapture system structure, TPS and their 

associated aero-thermal design basis are described in 

more detail by Justus, et al5

 (Titan atmosphere), 

Masciarelli, et al6

 (Guidance algorithms), Way, et 

al7

(Simulation), Takashima, et al8

(Aerothermodynamics), Oleiniczak, et al9

 (Radiative 

heating), and Laub10 (TPS). 

The study team started with the largest possible 

aeroshell, 3.75m diameter, that would fit inside a 4m 

launch vehicle fairing in anticipation of needing a large 

diameter high gain antenna. As discussed earlier, 

changing to a deployable HGA would allow a smaller 

diameter aeroshell. The selection of the 70 degree cone 

angle is discussed by Masciarelli, et al6

. 

Telecom 

The Orbiter telecom system supports two primary links, 

X-Band to Earth and UHF to the AP. Table 6 

summarizes the driving data return links. The X-Band 

System utilizes SDSTs and 50W TWTAs for 

communicating to Earth. The UHF telecom system is 

based on a next generation Electra Radio. 

Table 6. Orbiter Telecom Links 

Mission 

Phase

Orbiter 

Antenna 

(dBi)

Ground 

Station

Dist 

(AU)

Data 

Rate 

(bps)

Xmit 

Pwr 

(W)

Launch -6, Patch 34m BWG 0.5 10 50

Cruise

24.8, 

Printed 

Dipole 70m 11 500 50

Aero 

capture

Orbit 

Science

44.3, 

Fixed Dish 70m 11 75000 50

Mission 

Phase

Orbiter 

Antenna 

(dBi)

Probe 

Antenna 

(dBi)

Dist 

(km)

Data 

Rate 

(bps)

Xmit 

Pwr 

(W)

Probe 

Entry / 

Descent

-3 @ 30o

, 

2x2 patch

-3 @ 60o, 

Omni 85000

carrier 

only 4

Science 

Relay 5, Helix

-3 @ 60o

, 

Omni 2050 64000 4

X-Band

UHF

None

Power 

At 10 AU, solar power was out of the question. A 

Mutli-Mission Radioisotopic Thermal Generator 

(MMRTG) unit was selected for the Orbiter power 

source. This unit is currently in development and 

should reach TRL 6 by 2006. The expected 

performance of the MMRTG is approximately 6.3% 

efficiency for a 2000W thermal input. 

The efficiency of the units is tied to, among other 

things, the temperature differential of the unit. The 

6.3% efficiency is related to a finned exterior radiating 

to space. Cooling the exterior of the unit below what is 

expected from the finned radiator design will yield 

better power output. There is a potential with the 

Orbiter’s passive loop heat pipes to cool the surface of 

the MMRTGs to improve the power efficiency. To raise 

the power margin listed in Table 4 from 25% to 30% 

requires a power conversion efficiency of only 6.7%. 
21 

The power system includes 2 MMRTGs, total power 

available of 252W BOL, 214W at end of mission. 

Secondary batteries are included to help during peak 

periods with a typical assortment of battery charge 

controllers, power switching, and power conversion 

electronics. 

ACS 

Because of the limited power, reaction wheels were 

discarded in favor of thrusters capable of 0.7 mN-s 

impulses. All other ACS components are fairly standard 

equipment including star trackers, IMUs, and 

propulsion driver electronics. The Orbiter does not 

possess sun sensors because there is no critical need to 

sun point during a spacecraft upset. The Orbiter 

possesses 2 star trackers, 2 optical navigation cameras, 

and 2 C&DH strings which should suffice to allow the 

spacecraft to determine its attitude and point the HGA 

at Earth instead of solar panels at the sun during an offnominal event. 

Propulsion 

The propulsion system is a blow down hydrazine 

monopropellant system with two sets of thrusters. The 

first set of thrusters is comprised of 12 MIT thrusters, 

each with a 0.7 mN-s minimum impulse capability to be 

used for fine attitude control. The MITs are currently in 

a flight qualification process on track for TRL 8 before 

2006. The second set is comprised of 12 MR-120B 

engines, each with a 133.5 N force capability. The MR120Bs are used for attitude control during aerocapture 

and for Titan orbit maneuvering. 

Six thrusters are put on a dedicated line with latch 

valve, for a total of 4 latch valves, to ensure single fault 

tolerance (in degraded performance mode) against loss 

of mission. The single tank is a 74cm diaphragm tank 

with a Titanium shell. 

C&DH 

The C&DH system is JPL X200013 based. The cards 

selected are currently planned for TRL-6 by 2006, but 

an MRO based C&DH system might provide a lower 

risk technology solution in a sufficiently low mass and 

power package. 

Thermal 

The mission design presents several challenges for the 

thermal design: 

1. The MMRTGs together generate over 3700W of 

thermal heat. 

2. The MMRTGs are enclosed in an aeroshell 

designed to keep heat from getting in. 

3. The radiator system has to be designed to work 

before, during, and after aerocapture 

4. Inside the aeroshell, the system will experience 

solar distance of 0.95 AU (0.7 for Venus Gravity 

Assist) to 10 AU. 

Because Venus gravity assists were considered in the 

mission trade space3

, a 0.7 AU minimum solar distance 

was assumed for the thermal design efforts. A ~30 node 

lumped mass model of the spacecraft was constructed 

to compute temperature distributions during the key 

mission phases for various design options. 

Titanium/Water loop heat pipes running to hot radiators 

mounted on the Orbiter / Probe truss were chosen to 

solve the problem of getting the heat out of the 

aeroshell, and these were found to work even in the 0.7 

AU hot case at Venus. Aluminum / Ammonia loop heat 

pipes were also added to transport Orbiter electronics 

heat out of the aeroshell. A second set of 

Titanium/Water LHP carry MMRTG heat directly to 

the hydrazine tank. 

The computational model results confirmed that all of 

the key avionics and propulsion components were 

maintained within prescribed operating temperatures 

during both the cruise to Saturn and after orbit insertion 

when the aeroshell was jettisoned and the orbiter 

exposed to the cold space environment at Titan. 

For the aerocapture phase, it was assumed that the 

radiators were separated from the aeroshell 40 minutes 

before aerocapture and that the aerocapture lasted 20 

minutes, a total of 60 minutes without radiators. The 

heat from the MMRTGs and from the high speed entry 

is simply absorbed by the thermal mass of the vehicle 

during this time. In the thermal analysis, the entry 

heating was approximated by an instantaneous jump to 

250o

C in the bondline temperature between the TPS and 

structure, which is a conservative assumption. The 

results demonstrated that all orbiter components were 

within their operational temperature ranges; although 

some with small margins. 

Structure 

The structure is discussed in more detail by Hrinda11. In 

general, the Orbiter primary structural design was 

driven by: 

1. LV frequencies for Orbiter and Probe. 

2. Combined geometry constraints of SEPM, Orbiter, 

and Probe in LV fairing. 

3. LV loads for Orbiter and Probe. 

NEW TECHNOLOGY DEVELOPMENT 

Of all the technologies proposed that are currently less 

than TRL 6, only the TPS materials require additional 

funding to test the materials against the radiative heat 

loads expected at Titan9

. The other technologies that are 
22 

not currently at TRL 6: MMRTG, SEP Engine, and 

SEPM solar arrays, are all currently funded to reach 

TRL 6 in the 2006 time frame. If none of these three 

technologies actually reach TRL 6 by 2006, then the 

mission could still be performed with the currently 

lesser (mass) efficient technologies and a larger launch 

vehicle. 

RECOMMENDED ADDITIONAL ANALYSIS 

Many questions and trades consistent with continued 

Phase A/B efforts were identified by the study team. A 

summary of these issues is presented below along with 

a general classification of the issue as a lien, or 

opportunity, or either. 

• Launch Vehicle: Verify 4m fairing not available 

for Delta 4450 (Opportunity). 

• Launch Configuration: Eliminate structure 

through primary TPS (Lien). 

• SEP Propulsion Module: 1) Incorporate latest 

Glenn Research SEP Engine capability (either). 2) 

Develop solar array deployment sequence concept 

and verify associated structures and mechanisms 

mass (Lien). 

• Atmospheric Probe: 1) Verify 400kg is adequate 

(either). 2) Develop separation plane concept 

between AP and Orbiter which handles AP spin 

eject and thermal issues (AP MMRTG radiators) 

for 30 day coast to entry interface (Lien). 

• Navigation: Verify use of MITs does not degrade 

navigation performance beyond mission 

requirements (Lien). 

• Science Instruments: Develop conceptual designs 

for Multi-Spectral Imager and Microwave 

Spectrometer and verify TRL, mass, power, 

volume estimates (Lien). 

• Power: 1) Develop detailed power modes and 

profiles (either). 2) Verify 2 MMRTGs are 

adequate for full mission (Lein). 3) Verify 

conversion efficiency of MMRTG based on 

thermal design (Opportunity). 3) Verify EMI/EMC 

compatibility for component configuration (either). 

• Thermal: Verify MMRTG heat can be effectively 

routed to hydrazine system (manifolds, lines, 

thrusters) to eliminate need for heaters (Lein). 

• Telecom: 1) Investigate trade between Ka-Band 

system or 6m deployable antenna for X-Band 

system (either). 2) Verify UHF line of sight for 

AP-Orbiter link during EDL and science relay are 

consistent with antennas and pointing concept 

(Lien). 3) Add LGA/MGA for Earth acquisition 

prior to high bandwidth links (Lein). 

• Aeroshell: 1) Optimize packaging for smaller 

aeroshell (Opportunity). 2) Verify heating and TPS 

for new ballistic coefficient (either). 

• Cost: Generate cost estimate for complete flight 

system (either). 

CONCLUSIONS 

The study demonstrates general technical feasibility for 

a Titan Explorer Orbiter flight system designed to use 

aerocapture as the Titan orbit insertion mechanism. 

Many liens exist against the conceptual design 

presented, but opportunities balancing the liens also 

exist. A change from the medium launch vehicle to a 

heavy lift launch vehicle would help retire many of the 

leins without invalidating the feasibility of the mission 

concept. Technology readiness for the flight system is 

good with all major components currently being funded 

to achieve TRL 6 by 2006 to support a possible launch 

date as early as 2010. 

ACKNOWLEDGEMENTS 

The spacecraft and mission design described herein 

were built upon a Titan aerocapture mission study 

performed by JPL’s Team-X12. Although the 

configuration presented here is very different from the 

Team-X design, the Team-X design provided a valuable 

reference point from which system wide trades could be 

evaluated more easily. 

Personnel from Langley Research Center, Johnson 

Space Center, Ames Research Center, Marshall Space 

Flight Center, and the Jet Propulsion 

Laboratory3,4,5,6,7,8,9,10,11 were instrumental in 

determining system level requirements and subsystem 

capabilities for trade and mission performance analysis. 

Discipline experts at the Jet Propulsion Laboratory 

including Dave Hansen (Telecom), Brian Okerlund 

(Computer Aided Design), Jonathan Lam (Structural 

Analysis), John Huang (Antennas), Ray Baker 

(Propulsion), Rolando Jordan (SAR Instrument), Nick 

Mardesich (Solar Arrays), Bill Nesmith (MMRTGs); 

Applied Sciences Laboratory including Chern-Jiin Lee 

(thermal); Swales Aerospace including Michael 

Nikitnin (Loop Heat Pipes); all provided valuable input 

into the subsystem conceptual designs represented in 

this paper. 

Funding for this work was provided by the NASA InSpace Propulsion Program, managed by Les Johnson, 

MSFC. 

BIBLIOGRAPHY 

1. Chyba, C. et al., “Report of the Prebiotic 

Chemistry in the Solar System CSWG”, NASA 

internal report,1999. 

2. Tamppari, L. et al., “Cassini/Huygens Follow-On 
23 

Mission Study Science Report”, JPL internal 

report, 2001. 

3. M. Noca, R.W. Bailey, C.J. Sauer, and R.E. Dyke, 

“Titan Explorer Mission Trades from the 

Perspective of Aerocapture”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4801, July 

22, 2003. 

4. R.J. Haw, “Titan Approach Navigation for the 

Titan Aerocapture Orbiter”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4802, July 

22, 2003. 

5. C. Justus, A. Duvall, D. Johnson, “Engineering 

Level Model Atmospheres for Titan and Neptune”, 

39th AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4803, July 

22, 2003. 

6. J. Masciarelli, E. Queen, “Guidance Algorithms for 

Aerocapture at Titan”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4804, July 

22, 2003. 

7. D. Way, R.W. Powell, J. Masciarelli, B. Starr, K.T. 

Edquist, “Aerocapture Simulation and Performance 

for the Titan Explorer Misson”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4951, July 

22, 2003. 

8. N. Takashima, B. Hollis, K. Sutton, M. Wright, J. 

Oleiniczak, “Preliminary Aerothermodynamics of 

Titan Aerocapture Aeroshell”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4952, July 

22, 2003. 

9. J. Oleiniczak, D. Prabhu, M. Wright, N. 

Takashima, B. Hollis, K. Sutton, “An Analysis of 

the Radiative Heating Environment for 

Aerocapture at Titan”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4953, July 

22, 2003. 

10. Laub, B., “Thermal Protection Concepts and Issues 

for Aerocapture at Titan”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4954, July 

22, 2003. 

11. G. Hrinda, “Structural Design of the Titan 

Aerocapture Mission”, 39th 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Paper Number AIAA-2003-4955, July 

22, 2003. 

12. T. Sweetser, et al., “Titan Lander Conservative 

Science 01-06”, Team-X Report, June 4,5,8,15, 

2001. 

13. “Europa Orbiter/X2000 Avionics Development 

Industry Briefing”, NASA/JPL, Pasadena, Ca, June 

6, 2001. 
24 

TITAN EXPLORER MISSION TRADES FROM THE PERSPECTIVE OF 

AEROCAPTURE 

Muriel Noca, Robert W. Bailey

Jet Propulsion Laboratory 

California Institute of Technology 

Pasadena, California 

Contact: Muriel.noca@jpl.nasa.gov

A detailed Titan aerocapture systems analysis and spacecraft design study was performed as part of NASA’s InSpace Propulsion Program. The primary objective was to engineer a point design based on blunt body aeroshell 

technology and quantitatively assess feasibility and performance. This paper reviews the launch vehicle, propulsion, 

and trajectory options to reach Titan in the 2010-2015 time frame using aerocapture and all-propulsive vehicles. It 

establishes the range of entry conditions that would be consistent with delivering a 360 kg entry vehicle plus a 580 

kg orbiter to Titan. Results show that inertial entry velocities in the range of 5.3 to 6.6 km/s are to be expected for 

chemical and solar electric propulsion options with Venus and/or Earth gravity assists. Trip times range from 

approximately 6 years for aerocapture orbiters to 8-11 years for all-propulsive vehicles. In addition to trip time 

reduction, the use of aerocapture enables the mission with a Delta 4450 class launch vehicle as opposed to an allpropulsive orbit insertion approach, which requires a Delta IV heavy or Titan IV class launch vehicle. 

INTRODUCTION 

As part of the NASA In-Space Propulsion Program, 

aerocapture was investigated as an option for orbit 

insertion around Titan, the largest Moon of Saturn. This 

study involved several NASA centers and had for 

objective to conceptually design an aerocapture system 

for a generic orbiter/lander mission. This paper 

provides an overview of the mission trades performed 

during this study. The main objectives of the mission 

trades were to: 

1. Identify potential mission architecture and 

trajectories for a launch circa 2010-2015, 

which meant to identify launch vehicle 

options, launch opportunities and sensitivities, 

and potential trajectories using chemical 

ballistic propulsion and solar electric 

propulsion (SEP); 

2. Understand the sensitivities in flight time and 

Titan atmosphere’s inertial entry velocities; 

3. Provide a baseline trajectory and mission 

timeline. 

The level of analysis for the mission trades varied from 

relatively very detailed, in the case of the aerocapture 

system and SEP trajectory optimization, to more 

parametric in the case of the chemical system. The 

approach was to survey as much as possible the 

trajectory trade space, both for chemical with multiple 

gravity assists and for SEP with a wide range of flight 

times and various gravity assist options. Once the 

trajectories were compiled, the delivered mass at Titan 

was calculated given the maximum performances of 

representative launch vehicles. This delivered mass was 

then compared to the actual mass needed for Titan 

orbiter and lander design, thus highlighting the benefits 

of aerocapture. 

This paper first briefly describes the Titan Explorer 

lander and orbiter and then summarizes the 

transportation architectures considered. It then 

describes the findings for chemical and SEP transit to 

Saturn options, system and trajectories. It also describes 

briefly the aerocapture system and the baseline 

trajectory. Finally, it shows the overall architecture 

trade results. 
25 

SPACECRAFT DESCRIPTION 

The science objectives and basic spacecraft concept of 

this Titan Explorer mission were based on previous 

studies performed internally at the Jet Propulsion 

Laboratory.1

 The mission includes a landed module and 

an orbiter.2

 The lander was considered here as a black 

box, and only the navigation aspects of carrying this 

lander were taken into account. The lander performs a 

direct entry, independently of the orbiter. The orbiter 

was designed to perform aerocapture and modified in 

the trades when a chemical insertion was performed 

instead. The baseline concept uses Solar Electric 

Propulsion (SEP) to reach Saturn/Titan. Figure 1 shows 

the launch configuration of the overall spacecraft. 

Figure 1: Titan Explorer spacecraft launch 

configuration. 

TRANSPORTATION ARCHITECTURES 

To understand the sensitivities in aerocapture entry 

conditions into Titan’s atmosphere, it was necessary to 

perform a trade study of the various and most probable 

transportation options to Titan’s orbit in the 2010-15 

launch time frame. The transportation options for 

launch and transit from Earth to Titan were the 

following: 

- Option 1: Ballistic with or without a chemical 

stage, launch to a high positive C3, with gravity 

assists. 

- Option 2: Solar Electric Propulsion, launch to a 

low positive C3, with gravity assists. 

The final science orbit around Titan was a 1700-km 

altitude circular orbit. Thus the orbit insertion options 

considered were: 

- Option 1: Chemical insertion. 

- Option 2: Aerocapture with a chemical burn for 

periapsis raise. 

All four combinations of transportation were evaluated 

and will be described. 

Aerobraking, which consists of low orbit insertion via 

several passes in a planetary atmosphere, was not 

considered at the time of the study. A more recent study 

on the possibility of performing aerobraking in Titan’s 

atmosphere3

 has shown that aerobraking at Titan is 

limited by the gravitational perturbations of Saturn. The 

apoapsis of the aerobraking orbit would have to be 

quite low (below the 10000 km altitude range) to be in 

a gravitationally stable orbit around Titan. Higher 

apoapsis altitudes feature large spread in periapsis 

altitude (800 km spread in periapsis at an apoapsis 

altitude of 16000 km), making it very difficult to plan 

for and maintain aerobraking orbits. With this 

restriction, the delta-V saving of aerobraking compared 

to a direct insertion is quite low (~100-200 m/s), which 

limits its benefits. 

CHEMICAL BALLISTIC TRAJECTORIES TO 

TITAN 

Earth to Saturn Trajectories 

Ballistic direct trajectories as well as gravity assist 

trajectories were computed and gathered by Jon Sims 

and Carl Sauer from the Jet Propulsion Laboratory 

Trajectory Group for a launch period between 2010 and 

2016. Table 1 summarizes the performances of these 

trajectories. The maximum launch injected mass could 

then be found given the launch C3 for a Delta 4450 and 

a Delta IV Heavy. These launch vehicles were picked 

as representative of a range of launch vehicle 

performance. The launch vehicle data was provided by 

the NASA KSC Launch Support Group4

 and a 10% 

margin was held against the KSC provided performance 

(consistent with the JPL Team X conceptual design 

guidelines at the time of the study). Also note that some 

of the direct and gravity assist trajectories require a 

deep space maneuver, which Delta-V can sometimes be 

significant. 
26 

Table 1: Ballistic trajectories to Saturn/Titan.* EGA: Earth gravity assist, VEEGA: Venus Earth Earth 

gravity assist, VVVGA: Triple Venus gravity assist, JGA: Jupiter gravity assist, EJGA: Earth Jupiter gravity 

assist, VEEJGA: Venus Earth Earth Jupiter gravity assist. 

Launch 

date

Saturn 

Arrival 

Date Flight time

Vinf @ 

Saturn

Approach 

declinatio

n

Vinf @ 

Titan

Ventry 

inertial

Deep 

Space 

DV Launch C3

Delta 4450 

injected 

mass (kg)

Delta IV 

Heavy

(yrs) (km/s) (rad) (km/s) (km/s) (km/s) (km2/s2) 10% margin included

Direct 2/4/16 11/29/20 4.8 6.26 -0.401 5.39 5.84 0 109.3 - -

1/16/14 2/14/20 6.1 5.25 -0.262 4.33 4.88 0 105.2 - -

1/7/12 1/22/20 8.04 5.48 -0.105 4.10 4.67 0 106.1 - -

1/3/11 12/20/19 8.96 5.82 0.017 4.23 4.78 0 107.2 - -

11/26/09 9/18/14 4.81 5.93 -0.105 4.36 4.90 0.513 108.7 - -

12/31/09 11/15/19 9.87 6.17 0.070 4.47 5.00 0 108.4 - -

EGA 1/3/13 4/6/21 8.25 5.86 -0.349 4.96 5.45 0.407 48 - 3119

3/3/13 4/27/20 7.15 7.3 -0.436 6.16 6.55 0.504 47.7 - 3141

12/28/13 7/3/21 7.51 5.67 -0.332 4.80 5.29 1.488 28 2250 4905

3/16/14 9/24/21 7.53 5.49 -0.140 4.16 4.72 1.693 26.3 - 5085

VEEGA 4/3/12 3/13/22 9.94 7.83 0.436 6.50 6.88 0 12.2 3150 6746

4/3/12 3/13/22 9.94 6 0.436 5.39 5.84 0.217 12.1 3159 6759

VVVGA 7/26/10 12/3/18 8.4 9 -0.445 7.34 7.67 1.68 10.9 3249 6912

8/1/10 11/19/19 9.3 6.83 -0.417 5.79 6.21 1.35 9.7 3339 7106

7/1/13 12/31/23 10.5 6.5 -0.415 5.59 6.02 1.29 13.6 3051 6593

6/1/15 5/27/25 10 6.36 -0.447 5.64 6.07 1.1 15.1 2943 6381

JGA 12/16/16 8/7/20 3.64 15.74 0.436 12.77 12.97 0 97.8 - -

12/16/16 8/6/22 5.64 7.28 0.436 6.15 6.54 0 82 - -

12/22/16 11/14/23 6.89 4.54 0.436 4.68 5.19 0 80.8 - -

12/23/16 7/19/28 11.57 2.48 0.436 3.98 4.57 0 76 - -

EJGA 1/23/15 9/17/22 7.65 6.97 0.436 5.96 6.36 0.6695 25.6 2250 5162

VEEJGA 4/15/12 12/22/20 8.69 11.36 0.436 9.09 9.36 0 10.5 3272 6966

* The trajectories highlighted in yellow are the ones used in the subsequent analysis.

The trajectories highlighted in Table 1 were chosen for 

the overall mission trade as they represented a set of 

good performance trajectories. Careful consideration of 

all parameters (deep space Delta-V, approach velocity 

(Vinf) at Titan, launch C3, etc...) was used to make the 

choice of this set of trajectories. 

The spread in entry velocity is depicted in Figure 2 for 

these trajectories. Thus for aerocapture purposes, most 

of the trajectories have inertial entry velocities between 

5 and 7 km/s. This information will help select a 

baseline inertial entry velocity for a detailed 

aerocapture design, as will be discussed later. 

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

8.0

8.5

9.0

6 7 8 9 10 11

Flight Time to Titan (yrs)

Inertial Entry velocity @ 1000 km 

 EGA 2013-14

 VEEGA 2012

 EJGA 2015

 VVVGA 2010-15

 Direct 2009-2016

Figure 2: Titan inertial entry velocities for set of 

direct and gravity assist ballistic trajectories. 

As can be seen from Figure 2, direct trajectories offer 

lower inertial entry velocities, since the time it takes to 

perform the gravity assist has to be made up for in the 

cruise to Saturn. Also, increasing the number of gravity 

assists (for instance from a single Earth gravity assist to 

a triple Venus gravity assist) increased flight time and 

launch mass, and if not, it then increased significantly 

the inertial entry velocity. 
27 

Chemical Insertion Delta-V assumptions 

The chemical insertion Delta-V into a 1700-km altitude 

circular orbit around Titan can be computed given the 

Titan hyperbolic velocity (Vinf) and the approach 

declination. However, this Delta-V does not take into 

account a possible tour design around Saturn to pump 

down the initial orbit around Saturn using its satellites 

and to reduce the approach velocity around Titan. For 

the three trajectories picked, Table 2 summarizes the 

insertion Delta-V. The assumption was made that this 

Delta-V could be reduced by a moon tour down to 3 

km/s with a flight time penalty of 1-2 years. No 

trajectory analysis has been done to verify this 

assumption, and further analysis needs to be done to 

confirm this estimate. 

Table 2: Titan insertion ∆V for sample trajectories. 

 Flight 

time 

(yrs) 

Vinf at 

Titan 

(km/s) 

Titan 

insertion 

∆V (km/s) 

EGA 2013 7.5 4.8 3.8 

VEEGA 2012 9.9 5.4 4.3 

VVVGA 2010 9.3 5.8 4.7 

Chemical Propulsion system assumptions 

To perform the chemical deep space maneuvers or 

insertion burns, a generic bi-propellant system was 

assumed. The dry mass for this system is summarized 

in Table 3. The specific impulse of the chemical system 

was assumed at 325 sec. In addition, 10% of the 

deterministic propellant mass was held as propellant 

contingency for maneuver clean-ups. 

Table 3: Chemical propulsion system mass 

breakdown (includes 30% contingency). 

Subsystem Growth Mass 

(kg) 

Not scaled with propellant mass: 

- Propulsion 

- Thermal 

- Telecom+electronics 

- Structure 

19.5 

16.5 

2.3 

261.5 

Scaled with propellant mass: 

- Tank 

- Tank structure 

- Thermal 

5% 

4% 

1% 

SEP TRAJECTORIES TO TITAN 

Early on, an extensive database of Venus gravity assists 

SEP trajectories on a Delta 4240 was built, as they 

clearly provided better delivered mass for equivalent 

flight times compared to direct (no gravity assist) SEP 

trajectories. These trajectories served the purpose of 

evaluating the sensitivities in launch date, number of 

thrusters, power levels and inertial entry velocities. 

Subsequently, a smaller set of Earth gravity assist and 

Venus gravity assist on the Delta 4450 and Delta IV 

Heavy was calculated and used. 

Carl Sauer (JPL) ran the SEP trajectory optimization 

code named SEPTOP for Solar Electric Propulsion 

Trajectory Optimization Program, which is based on the 

calculus of variations. This code optimizes two body 

interplanetary trajectories and can model discrete 

numbers of operating Xenon thrusters throughout the 

trajectory. Carl allowed for a coast time duty cycle of 

10% to simulate times when the spacecraft is not 

thrusting due to housekeeping activities, and assumed a 

constant 250 W from the solar arrays for the spacecraft. 

Solar Electric Propulsion system assumptions 

The ion thruster used to calculate the SEP trajectories is 

and advanced 5-kW 5000 sec version of the flown 

NSTAR engine. The characteristics of the NSTAR 

technology can be found in many references.5,6 The 

description of the 5-kW derivative of NSTAR named 

NGN for “Next Generation NSTAR” can be found in 

reference [7, 8]. This thruster is characterized by 

differences in four major parameters compared to 

NSTAR: engine input power, maximum specific 

impulse, and engine total impulse (or throughput) 

capability. Table 4 shows the projected performances of 

NGN. 

Table 4: High-level NGN versus NSTAR 

characteristics. 

 NSTAR NGN 

Max. thruster processed 

power (kW) 

2.3 5 

Engine diameter (cm) 30 30 

Maximum Isp (sec) 3100 5000 

Xe throughput (kg) 130 250 

The ion propulsion system (IPS) was designed more as a 

propulsion module than just thrusters and power 

processing units. Figure 3 shows a simplified block 

diagram of the NSTAR IPS (single string). To that basic 
28 

configuration was added redundancy, structural and 

thermal considerations. Figure 3 also shows an example 

of what the IPS module designed here could look like. 

The solar arrays were sized based on a projection of the 

AEC-Able Ultraflex array capability. Since this array 

technology scales with power from ~ 1 kW up to ~ 30 

kW, it was used as a representative potential technology 

for SEP applications. The specific mass was assumed to 

be 200 W/kg at 24 kW. A 14% degradation factor was 

applied to the array Beginning-of-Life (BOL) power to 

account for various degradation phenomena. Also, in 

order to support power demand during launch, a 

primary battery was used prior to solar array 

deployment. 

The number of thrusters and PPUs was calculated on the 

basis of power requirements (4 minimum plus 1 

redundant) and thruster propellant throughput. The 

system architecture followed a conventional approach 

with parallel strings of PPUs and thrusters. Each PPU 

drives one thruster but is cross-strapped to two engines. 

One spare ion engine, one spare PPU and DCIU were 

also included for single-fault tolerance. Each thruster 

was gimbaled separately. The PPUs were assumed to be 

95% efficient. 

Solar Array

Propellant

Storage

Gimbal

Power

Conditioning Exhaust

S/C

C&D

Propellant

Distribution

and

Control

Digital

Interface &

Control Unit

Note: 

Structural and Thermal

Hardware Not Shown

Thruster

Solar Array Solar Array

Propellant

Storage

Gimbal

Power

Conditioning Exhaust

S/C

C&D

Propellant

Distribution

and

Control

Digital

Interface &

Control Unit

Note: 

Structural and Thermal

Hardware Not Shown

Thruster

Solar Array

Figure 3: Ion propulsion module block diagram and conceptual configuration for system sizing. 

The tank mass fraction was assumed to be 3.5% for 

Xenon when stored as a supercritical gas (~2000 psia). 

Furthermore, a 10% propellant contingency was added 

to the deterministic propellant mass to account for flow 

rate characterization, residuals, attitude control and 

margin. 

Since the system masses are function of mainly power 

level, launch mass and propellant mass, each trajectory 

was uniquely considered and had a system mass 

associated with it. The component and subsystem sizing 

assumptions are given in Table 5. To be consistent with 

the JPL Team X conceptual design guidelines at the time 

of the study, 30% mass contingency was applied to all 

spacecraft subsystems, and a 10% launch vehicle margin 

was assumed. 

Table 5: Ion propulsion system mass breakdown 

(includes 30% contingency). 

Subsystem Growth Mass 

(kg) 

Not scaled with propellant mass: 

- Propulsion 

- Power 

- Thermal 

- Telecom+ACS +electronics 

- Structure 

168.5 

192.8 

59.2 

7.6 

340.5 

Scaled with propellant mass: 

- Tank 

- Tank structure 

- Thermal 

3.5% 

4% 

1% 
29 

With the appropriate thruster model, trajectories were 

run for power level of 24 kW. Results are in terms of 

net delivered mass. The net delivered mass is defined as 

the spacecraft dry mass minus the dry mass of the ion 

propulsion system. Therefore the net delivered mass is 

everything on the spacecraft that isn’t propellant or part 

of the ion propulsion module. 

Net delivered mass and inertial entry velocity 

sensitivities to launch date, arrival date, SEP power 

level and thruster technology 

Figure 4 shows results of net delivered mass as a 

function of flight time and launch years for the NGN 

thruster and for the NSTAR thruster (for reference). 

These trajectories were run for a Delta 4240 launch 

vehicle. The desired net delivered mass for the 

spacecraft was on the order of 2000 kg. 

800

900

1000

1100

1200

1300

1400

1500

1600

1700

1800

4 5 6 7 8 9 10

Flight time (yrs)

Net delivered mass (kg)

NSTAR DS-1, 2008 launch

2008 launch

2013 launch

2015 launch

2010 launch

VGA Titan Explorer

Delta 4240, SEP Power 1AU is 24 kW

Next Generation NSTAR

Figure 4: Venus Gravity Assist SEP trajectories to 

Saturn as a function of launch date. 

Figure 5 shows the corresponding inertial entry 

velocities. As can be seen, the inertial entry velocity 

increases significantly for flight times below 5 years. It 

is also very dependent on flight time, launch date, and 

thruster technology. However, as Figure 6 shows, the 

inertial entry velocity was only weakly dependent on 

SEP power for a given launch date and thruster 

technology. There is also a significant variability in the 

choice of gravity assist. Thus choosing a flight time 

range will determine a range of inertial entry velocities. 

Flight times around 6 to 7 year offer the most “net 

delivered mass” benefit and result in entry velocities 

less than 7 km/s for most launch opportunities. 

4

5

6

7

8

9

10

4 5 6 7 8 9 10

Flight time (yrs)

Inertial entry velocity @ 1000 km alt. (km/s).

2008 launch NSTAR DS-1, 2008 launch

2013 launch

2015 launch

2010 launch

VGA Titan Explorer

Delta 4240, SEP Power 1AU is 24 kW

Next Generation NSTAR

Figure 5: Titan inertial entry velocity at 1000 km 

altitude as a function of flight time. 

The weak sensitivity to the SEP power level is mostly 

due to the fact that over the range of power looked at, 

the trajectory optimization code is going to try to follow 

the same optimum acceleration path. Thus for high 

power level, it will optimize the trajectory at lower 

launch C3, thus injecting more mass. The acceleration, 

which is proportional to the power level to mass ratio 

will be roughly the same as a low power, large C3, low 

launch mass case. Since it will follow almost the same 

trajectory profile, the arrival hyperbolic velocity will 

only vary slightly (such variation could be seen by 

zooming in Figure 6). This is true for a fixed flight time 

and launch date. 

4

5

6

7

8

9

10

10 12 14 16 18 20 22 24 26 28 30

SEP power (kW, 1AU EOL)

Inertial Entry Velocity @ 1000 km alt. (km/s).

VGA Titan Explorer

Delta 4240, Optimum flight time 7.75 yrs

Launch date 2010

NSTAR DS1

NGN

Figure 6: Inertial entry velocity at 1000 km altitude 

for Venus Gravity Assist SEP trajectories to Saturn 

with Delta 4240, fixed flight time (7.75 years) but 

varying SEP power level. 

The launch window to perform a given gravity assist 

(VGA or EGA) is about one month. The sensitivity in 

propellant mass for that window is included in the 10% 

propellant margin. 

Careful consideration was given to the sensitivity in 

arrival date. Since Titan’s orbital period around Saturn 
30 

is 16 days, the orbit geometry varies significantly 

depending on the arrival day. Figure 7 illustrates that 

point. However, it is possible to tune the arrival date 

with the SEP system or with a small chemical Delta-V 

at the end of the SEP phase to target a desired entry 

condition. The pattern shown in Figure 7 repeats every 

16 days. 

Figure 7: Sensitivity of Titan hyperbolic excess 

velocity as a function of arrival date. 

Net delivered mass and inertial entry velocity 

sensitivities to gravity assist and launch vehicle 

After looking at sensitivities in launch dates, arrival 

dates, SEP power level and thruster technology for a 

given gravity assist type and launch vehicle, the study 

called for more trajectories to perform the overall trade 

study. The SEP trajectories presented here assume a 

2010 launch date and 24-kW SEP system with 4 

operating NGN thrusters. The launch vehicle selected 

were the Delta 4450 and Delta IV Heavy, to enable 

more mass to be delivered, for both a Venus and Earth 

gravity assist. Figure 8 shows the net delivered mass for 

all 4 trade options. Figure 9 shows the corresponding 

inertial entry velocities. Also added to these figures are 

the points selected for the trade study. 

Selection of the inertial entry velocity 

In view of these results, it was decided that an entry 

velocity of 6.5 km/s would represent the best 

compromise between short flight times and high net 

delivered masses. Although somewhat arbitrary, it was 

felt that the aerocapture design would not change 

significantly for inertial entry velocities between 6 and 

7 km/s. Using inertial entry velocities below 5.5 km/s 

or so would probably be feasible but would reduce the 

aerocapture performance and robustness. 
31 

500

1000

1500

2000

2500

3000

3500

4000

4500

5000

4 5 6 7 8 9 10

Flight time (yrs)

Net delivered mass (kg)

VGA Delta 4450

VGA Delta IV Heavy

EGA Delta 4450

EGA Delta IV Heavy

Titan Explorer

SEP Power 1AU is 24 kW

Launch date 2010

Next Generation NSTAR

Trajectories 

picked for 

trade study

Figure 8: Venus and Earth gravity assist SEP trajectories to Saturn as a function of launch vehicle. 

4

5

6

7

8

9

10

4 5 6 7 8 9 10

Flight time (yrs)

Inertial Entry Velocity @ 1000 km alt. (km/s).

VGA Delta 4450

VGA Delta IV Heavy

EGA Delta 4450

EGA Delta IV Heavy

Titan Explorer

SEP Power 1AU is 24 kW

Launch date 2010

Next Generation NSTAR

Trajectories 

picked for 

trade study

Figure 9: Titan inertial entry velocity at 1000 km altitude as a function of flight time. 
32 

Baseline trajectory 

The baseline trajectory for the design of the aeroshell 

and other components of the aerocapture system was 

selected based on the following criteria: 

- The mission architecture should use the smallest 

launch vehicle possible to reduce cost; 

- The trajectory performance should provide 

adequate system mass margin (30%) for growth, 

and adequate system reserves (> 10%); 

- The trajectory should provide a Titan inertial 

entry velocity close to 6.5 km/s consistent with 

the design of the aeroshell. 

Thus the trajectory selected featured (see Figure xx): 

- Launch vehicle: Delta 4450 (5 m fairing) 

- Flight time: 5.9 years 

- Launch date: 12/24/2010 

- Arrival date: 11/17/2016 

- Earth Gravity Assist: 12/03/2012 

- Launch C3: 8.6 km2/s2 

- Launch mass: 3423 kg 

- Propellant mass: 460 kg deterministic 

- Vhyp @ Saturn: 7.97 km/s 

- Ventryinertial: 6.5 km/s @ 1000 km 

- Thrusters: 4 operating NGN 

- SEP power level: 24 kW 

A detailed approach and navigation analysis was 

performed9

 and suggested the timeline shown in Figure 

11 for this baseline trajectory. This timeline includes 

Trajectory Correction Maneuvers (TCM) for both the 

lander and the orbiter targeting. 

012 3456789

Time in years

Thrusting cruise Coast Science

SEP module release

Entry day targeting TCM

Launch

All check-out

Earth Gravity Assist

Time in days

-60 -50 -40 -30 -20 -10 0

TCM 1

Cruise error

corrections

Probe targeting

Probe release

TCM 2

Probe 

targeting

TCM 3

Orbiter 

targeting

TCM 4

Orbiter 

cruise 

error

corrections

TCM 5

Final orbiter 

targeting

Time in hours

-6 -4 -2 -0 2 4 6

Atm. Entry

TCM 6

Optional Aerocapture

Heatshield

release

Periapsis

Raise 

Maneuver

012 3456789

Time in years

Thrusting cruise Coast Science

SEP module release

Entry day targeting TCM

Launch

All check-out

Earth Gravity Assist

Time in days

-60 -50 -40 -30 -20 -10 0

TCM 1

Cruise error

corrections

Probe targeting

Probe release

TCM 2

Probe 

targeting

TCM 3

Orbiter 

targeting

TCM 4

Orbiter 

cruise 

error

corrections

TCM 5

Final orbiter 

targeting

Time in hours

-6 -4 -2 -0 2 4 6

Atm. Entry

TCM 6

Optional Aerocapture

Heatshield

release

Periapsis

Raise 

Maneuver

Figure 11: Baseline trajectory timeline.

AEROCAPTURE SYSTEM 

The aerocapture system is described in detail in 

reference [2] and [10]. It has been designed for the 

baseline trajectory, which lead to an inertial entry 

velocity of 6.5 km/s. 

Figure 10: Baseline SEP trajectories to Saturn. 

The heatshield design was based on a 70 deg. half coneangle using the Viking-Pathfinder heritage, and was 

sized to fit the orbiter 2.4 meter diameter high gain 

antenna.. Table 6 summarizes the mass breakdown for 

the aeroshell system. As can be seen, the total dry mass 

of the aeroshell system is 426 kg for a total entry dry 

mass of 1026 kg (~ 41.5% aeroshell entry dry mass 

fraction). Other aerocapture-related hardware was 

ejected before entry. This hardware is also summarized 

in Table 6. 

Table 6: Aerocapture system mass breakdown 

(includes 30% contingency). 

Subsystem Growth Mass 

(kg) 

Mass that entered the atmosphere: 

- Heatshield, backshell and 

structure 

- Hydrazine propellant 

426 

89 

Aerocapture mass jettisoned prior 

entry: 

- ACS, telecom, thermal 

radiators and loop heat pipes 

for the RTG, intruments 

72 
33 

MISSION ARCHITECTURE TRADE RESULTS 

The overall mission architecture trade results are 

summarized in Table 7. This table shows first the type 

of launch vehicle followed by the gravity assist type, 

the transit propulsion system and the Titan capture 

system. It assumes that the full capability of the launch 

vehicle is used and calculates the payload surplus or 

deficit mass compared to the mass required at Saturn 

before insertion. 

Table 7 also shows additional structure mass not part of 

the aerocapture system. The Pre-insertion ejected mass 

includes about 62 kg for the orbiter to lander interface 

structure and about 71 kg for ACS, telecom, thermal 

radiators and loop heat pipes for the MMR RTG and 

other. In the case of chemical insertion, the orbiter to 

lander interface is assumed not to be jettisoned and thus 

is included in the Payload in Titan orbit mass. The 

detailed mass breakdown can be found in [2]. Figures 

12 and 13 render some of the results of Table 7. They 

show the payload surplus or deficit mass as a function 

of transit propulsion, gravity assist and launch vehicle 

for chemical or aerocapture insertion. The payload 

surplus or deficit mass is the mass above or below the 

necessary mass to deliver the lander and orbiter around 

Titan. It does not include the lander or orbiter mass. 

Both figures clearly show the advantages of 

aerocapture, which in every case looked at provided 

more payload reserve and shorter flight times than for a 

chemical insertion burn. However, they also show that 

it is possible to deliver sufficient payload mass (low 

margin) with an all chemical insertion system. Here 

again, the penalty will be flight time. 

CONCLUSIONS 

This paper summarizes the transit trajectory options for 

the Titan Explorer and derives the range of entry 

conditions for the aerocapture maneuver inside Titan’s 

atmosphere. This survey shows that inertial entry 

velocities in the range of 5.5–7 km/s are to be expected. 

This range offers the best combination of highest 

delivered mass to Titan’s orbit and lowest entry 

heating. The study chose to baseline an inertial entry 

velocity of 6.5 km/s for the detailed design of the 

aerocapture system, and the corresponding SEP 

trajectory is provided. 

-2000

-1000

0

1000

2000

3000

4000

5000

VVVGA 

D4H Chem

VVVGA 

D4H Aero

VEEGA 

D4H Chem

VEEGA 

D4H Aero

EGA 

D4H Chem

EGA 

D4H Aero

Payload surplus or deficit (kg)

4

5

6

7

8

9

10

11

12

Flight time (yrs)

Payload

Flight time

Transit Propulsion: Chemical Ballistic

Figure 12: Payload surplus or deficit mass for the 

chemical ballistic transit cases for the Delta IV 

Heavy.

-1000

-500

0

500

1000

1500

2000

2500

3000

3500

EGA D4450

Chem

EGA D4450

Aero

VGA D4H

Chem

VGA D4H

Aero

EGA D4H

Chem

EGA D4H

Aero

Payload surplus or deficit (kg)

4

5

6

7

8

9

10

11

12

Flight time (yrs)

Payload

Flight time

Transit Propulsion: SEP

Figure 13: Payload surplus or deficit mass for the 

SEP transit cases for the Delta 4450 and Delta IV 

Heavy.

This paper also summarizes the mission transportation 

trades performed during the study to show the benefits 

of aerocapture. The study shows that aerocapture as a 

norbit insertion option provides more delivered mass in 

every launch vehicle and gravity assist case looked at 

than chemical insertion, and shorter flight time 

(typically by 2 years). However, all chemical or SEP 

with chemical insertion cases exist that would deliver 

the necessary mass in Titan’s orbit with a Delta IV 

Heavy with flight times between 8–11 years. The 

baseline trajectory case for this study is an SEP 

aerocapture case on a Delta IV 4450 with a flight time 

of 5.9 years. 
34 

Table 7: Overall mission architecture trade results. 

Launch Vehicle

Gravity Assist VVVGA VGA VEEGA VEEGA

Earth to Saturn Prop System Chem SEP Chem SEP SEP Chem Chem

Titan Capture Aero Aero Aero Chem Aero Chem Aero Chem Chem Chem Aero Chem Aero Chem Aero

7

Cruise Time to Titan (yrs) 9.3 5.7 7.5 ~7.8 5.9 ~11.3 9.3 9.9 ~11 ~7.7 5.7 ~8.5 7.5 ~8.1 6.1

Launch C3 (km2/sec2) 9.7 10.2 28 8.4 8.6 9.7 9.7 12.1 12.1 36.2 36.2 28 28 17.8 17.8

SEP Power (kW) --- 24 --- --- 24 --- --- --- --- 24 24 --- --- 24 24

Entry Velocity (km/s) 6.2 6.4 5.3 5.3 6.5 6.2 6.2 5.9 5.9 6.4 6.4 5.3 5.3 6.5 6.4

Earth to Saturn Cruise ∆V (km/s) 1.35 8.3 1.49 6.4 6.7 1.35 1.35 0.2 0.2 6.9 6.9 1.49 1.49 4.6 4.6

Titan Insertion Chem ∆V (km/s) 0 0 0 3 0 3 0 4.3 3 3 0 3 0 3 0

3 5 5 5,6 5 5,6

Launch Capability 3339 3298 2250 3439 3423 7106 7106 6759 6759 4135 4135 3141 5450 6019 6019

Propellant Mass1 1239 619 893 1485 506 5464 2671 5160 4363 2015 715 2253 2208 3257 633

LV to Prop Module Adapter 78 78 78 78 78 78 78 78 78 78 78 78 78 78 78

Prop Module Dry Mass 424 820 389 808 810 846 567 816 736 1257 828 525 521 1383 821

Available Payload Mass to Saturn 1598 1781 890 1068 2029 718 3790 705 1582 784 2514 285 2644 1301 4488

Prop Mod to Payload Adapter 61 61 61 61 61 61 61 61 61 61 61 61 61 61 61

Direct Entry Payload Mass (Lander) 364 364 364 364 364 364 364 364 364 364 364 364 364 364 364

Payload in Titan Orbit (Orbiter) 511 511 511 573 511 573 511 573 573 573 511 573 511 573 511

Aerocapture System2,4 515 515 515 515 515 515 515 515

Pre-Insertion Ejected Mass 133 133 133 73 133 73 133 125 73 73 133 73 133 73 133

Required Mass to Saturn 1584 1584 1584 1071 1584 1071 1584 1123 1071 1071 1584 1071 1584 1071 1584

Payload Surplus (Deficit) Mass 14 197 (694) (3) 445 (353) 2206 (418) 511 (287) 930 (786) 1060 230 2904

System Mass Margin 29.8% (LV - CBE) / LV

System Reserve 13% (LV - Growth) / LV

Assumptions and Notes:

7 Titan Aerocapture Study Reference Mission

EGA

Chem SEP

Delta IV Heavy

VVVGA VGA

Delta 4450

EGA

All masses are growth mass listed in kg

1

 Propellant mass calculated using "Launch Capability" as system total mass; Isp = 325, includes 10% mass contingency

2

 Aerocapture mass for Chemical Earth to Titan Prop Modules may change slightly (entry velocity not equal to 6.5 km/s)

Chem SEP

3

 This launch capability is extrapolated data

4 Aerocapture mass includes propellant for circularization delta-V

5

 Assumes delta V required for capture can be reduced to 3 km/s through Saturn/moon tour with flight time penalty. No supporting analysis 6

 Propellant mass and Prop Module Dry Mass for SEP / Chem options includes propellant and dry mass for both SEP and chemical stages

ACKNOWLEDGMENTS 

The research described in this paper was carried out at 

the Jet Propulsion Laboratory, California Institute of 

Technology, under a contract with the National 

Aeronautics and Space Administration. 

The authors gratefully acknowledge the contributions of 

Carl Sauer, Jon Sims, and Theresa Debban for all their 

trajectory creativity and calculations. The authors also 

wish to thank Eric Dyke (Swales Aerospace) and MaryKae Lockwood (LaRC) for their conscientious review 

of the results, John Brophy (JPL) for all the SEP 

thruster details, Ray Baker (JPL) for all the chemical 

system details, and the In-Space Propulsion Program 

for making the funding available for this work. 

REFERENCES 

[1] T. Sweetser et al., “Titan Lander Conservative 

Science 01-06”, Team X report, June 4-15, 2001. 

[2] R. Bailey, J. Hall, T. Spilker, “Titan Aerocapture 

Mission and Spacecraft Design Overview”, AIAA2003-4800, 39th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference, Huntsville, AL, July 2003. 

[3] Communication with Greg Whiffen, JPL, during 

the In-Space Propulsion SAIC Titan Explorer 

review, April 2003. 

[4] http://elvperf.ksc.nasa.gov/elvMap/

[5] Rayman, M. D., “Results from the Deep Space 1 

Technology Validation Mission,” IAA-99-

IAA.11.2.01, 50th International Astronautical 

Congress, 4-8 Oct 1999, Amsterdam, The 

Netherlands. 

[6] Polk, J. E., et al., “In-Flight Performance of the 

NSTAR Ion Thruster Technology on the Deep 

Space One Mission,” AIAA-99-2274, 35th

AIAA/ASME/SAE/ASII Joint Propulsion 

Conference, June 21-23, 1999. 

[7] Brophy, J.R, “Ion Propulsion System Design 

for the Comet Nucleus Sample Return Mission,” 

AIAA-2000-3414, 36th AIAA/ASME/SAE/ASII 

Joint Propulsion Conference, July 2000. 

[8] M Noca, “Next Generation Ion Engines: Mission 

Performances”, IEPC-03-313, 28th International 

Electric Propulsion Conference, Toulouse, France, 

March 17-21, 2003. 

[9] R. Haw, “Approach Navigation for a titan 

Aerocapture Orbiter”, AIAA-2003-4802, 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Huntsville, AL, July 2003. 

[10] M. Lockwood, “Titan Aerocapture Systems 

Analysis”, AIAA-2003-4799, 39th AIAA/ASME/ 

SAE/ASEE Joint Propulsion Conference, 

Huntsville, AL, July 2003. 
35

APPROACH NAVIGATION FOR A TITAN AEROCAPTURE ORBITER 

Robert J. Haw 

 AIAA Member 

Navigation and Mission Design 

Jet Propulsion Laboratory, California Institute of Technology 

Pasadena, CA 91109-8099 

robert.haw@jpl.nasa.gov

ABSTRACT 

A proposed mission will send an orbiter and surface probe to Titan. Aerocapture technology will slow the 

spacecraft at Titan and perform the orbit insertion. The navigation strategy uses ∆VLBI measurements and optical 

imaging in order to satisfy the flight system performance requirements. The performance requirement metric is the 

spacecraft’s atmosphere entry flight path angle, which must fall within -36.8° ±1.0 (3σ) (the error bar is 

preliminary). The requirement can be satisfied with a data cutoff 7 days before Titan, assuming a Cassini-era Titan 

ephemeris. There is margin in the arrival template to tighten (reduce) the entry corridor by scheduling a data cutoff 

closer to Titan, and a data cutoff 2 days before Titan is baselined here. Improvements to the performance are 

discussed by anticipating enhancements to the current level of technological readiness. The surface probe can 

satisfy an entry flight path angle requirement of -50°±5 (3σ) by separating any time within ~3 months of Titan. 

INTRODUCTION 

The Titan orbiter mission will utilize solar electric 

propulsion during interplanetary cruise to deliver an 

orbiter and probe to Titan, and will employ 

aerocapture technology to assist with orbit insertion. 

The work presented here is part of a Titan mission 

study described in an overview paper given in 

Reference 1. 

The focus of this paper is a demonstration of the 

feasibility of direct-entry aerocapture as a 

replacement for an orbit insertion maneuver. The 

probe is considered secondarily only in so far as 

illustrating that probe requirements can be met. 

This paper examines the navigation accuracies of the 

orbiter and probe as they approach and encounter 

Titan. Since there is a limit on the accuracy with 

which the initial state and subsequent dynamics are 

known, the future state cannot be computed with 

certainty from the initial one. The ‘delivery’ is 

defined as the uncertainty expected in the future 

spacecraft state (at its time-of-arrival) computed at 

time T (where T is before the time-of-arrival). That 

is, a Titan–2 day delivery represents the prediction 

(with dispersions) of the location of the spacecraft at 

arrival, when 2 days away from Titan. 

The epoch for this analysis is Titan-75 days. 

SPACECRAFT CONFIGURATION 

The design of the flight system is discussed in 

Reference 2. Externally, the flight system during 

Titan approach consists of two stacked aeroshells 

connected via an external truss. (The SEP propulsion 

module and associated solar array fans have been 

jettisoned earlier.) The spacecraft configuration is 

shown in Figure 1. The probe aeroshell is located on 

top and the orbiter aeroshell beneath. The probe and 

orbiter respectively are enclosed by the aeroshells, 

which consist of a backshell/ heatshield pairing. The 

aeroshell protects the contents from the high heat 

loads experienced during atmospheric entry. 

An external truss connects the two aeroshells. 

Mounted on the truss is a rear-looking medium gain 

antenna (MGA) for telecommunications and 

navigation. Also mounted on the truss are forwardlooking cameras for optical navigation. The cameras 

should have independent pointing control. 
36

The mass of the flight system (including propellant 

load) is 1465 kg. The orbiter represents 75% of the 

mass allocation and the probe/truss the remaining 

25%. 

ATTITUDE MAINTENANCE 

The spacecraft is assumed to be 3-axis stabilized with 

a momentum-wheel ACS. The momentum wheels 

maintain spacecraft pointing. Balanced 0.7 N 

minimum impulse thrusters perform scheduled 

momentum de-saturation burns. (Note: momentumwheel stabilization is not consistent with the baseline 

spacecraft design in Reference 2, which assumes 

limit-cycling ACS. The inconsistency has been 

identified but insufficient time was available to 

address this issue before the study ended.) 

SEP MODULE 

The SEP module consists of five Glenn Research 

Center 5 kW ion engines. Four engines operate at a 

time, leaving one engine in reserve. Each engine 

produces 0.15 N of thrust. Four solar array fans with 

a total area of 23 m2

 can generate 24 kWe at 1 a.u. 

[3]. The engines thrust while within the inner solar 

system, but when the solar range exceeds ~4 a.u. the 

SEP stage (including solar arrays) is discarded. 

PROPULSION 

The propulsion system after the SEP is jettisoned is a 

monopropellant hydrazine system. This subsystem 

must perform spin control, attitude control, and 

trajectory correction maneuvers (TCMs) during the 

approach to Titan. 

Thrusters are used to de-saturate momentum wheels 

and to correct trajectory errors. All propulsive 

functions are performed with the RCS thrusters 

acting through ports in the backshell of the orbiter 

aeroshell (ports are not shown in Figure 1). After 

orbit insertion (without the aeroshell) the RCS 

thrusters perform the Titan pericenter raise maneuver. 

TELECOMMUNICATIONS 

The telecommunications subsystem operates an 

X-band system for direct-to-Earth communications 

and a UHF system for communications with the 

probe during entry, descent and deployment (EDD). 

A truss-mounted X-band medium-gain antenna 

(MGA) (with a low gain antenna back-up) is the only 

communication link during interplanetary flight. The 

MGA design is a 24 dBi phase array antenna 28 cm 

in diameter. At 10 a.u. from Earth the data rate to a 

70 m ground antenna is 600 bps [2]. The probe 

backshell includes an aft-mounted LGA for UHF 

communications during EDD. 

Tracking and telemetry will use the Small Deep 

Space Transponder, which supports phase coherent 

two-way doppler and ranging, command signal 

demodulation and detection, telemetry coding and 

modulation, and differential one-way range (DOR) 

tone generation (tone sidebands at ±19 MHz). DOR 

tones are used for interferometric ∆DOR 

measurements. 

TRAJECTORY DESIGN 

The interplanetary reference trajectory characteristics 

are described in Reference 2. Since the reference 

trajectory has some fluidity during advanced study 

exercises (and changes frequently), the launch date 

chosen for this navigation investigation is December 

15, 2010, and may not agree with the latest reference 

trajectory. The launch mass is 2515 kg, including the 

SEP module. The interplanetary trajectory includes a 

gravity assist flyby of Venus in May 2012 and arrives 

at Titan on August 21, 2016. The hyperbolic excess 

velocity at arrival is 6.1 km/s and the range to Earth is 

9.8 a.u. (one-way light time equals 81 minutes). 

The orbiter and probe initially approach Titan on an 

impact trajectory. After the probe separates from the 

orbiter bus it follows a ballistic path into Titan. The 

probe is nominally targeted to a mid-latitude region 

in the northern hemisphere (target TBD). 

Before probe release the spacecraft is spun-up to a 

rate TBD (on the order of 3 rpm). After probe 
37

release the remaining orbiter bus and truss are spundown and a trajectory deflection maneuver (TDM) is 

performed. See Figure 2, where on the scale of that 

figure the probe release and TDM appear to occur 

simultaneously. The truss is jettisoned before 

aerocapture. 

The TDM serves two purposes: re-directing the 

orbiter’s trajectory to intercept the orbiter entry 

interface point (a lateral movement of approximately 

600 km in the B-plane), and slowing the orbiter with 

respect to the probe (a delay on the order of hours). 

This delay permits the orbiter to function as a 

communication relay for the probe during EDD. 

Figure 2: Entry Interface for Probe and Orbiter 

Orbiter delay is a trade-off between probe transmitter 

power and the duration of EDD and initial surface 

operations (to be relayed and recorded by the orbiter 

before aerocapture). In Figure 2 the blue line 

indicates the orbiter and the red line the entry probe. 

The position of the orbiter at the time of probe entry 

is indicated by the blue dot with the label “EI: probe” 

(Entry Interface for probe). For a delay of 3.5 hours, 

when the probe reaches its entry interface, the 

distance between probe and orbiter is approximately 

80,000 km. The orbiter subsequently enters Titan’s 

atmosphere at the blue dot labeled “EI: orbiter”. By 

this time the probe has been transmitting data to the 

orbiter for about 3.5 hours. 

The magnitude of the TDM as a function of probe 

release time and slow-down maneuver is 

parameterized in Figure 3. A probe release at E-30 

days and TDM at E-29 days represents the baseline 

unless otherwise stated. 

0

50

100

150

200

250

2 hr delay

3.5 hr delay

Magnitude (m/s)

Days Before Entry

-30 -20 -10 0

Figure 3: TDM magnitude v. probe release time

The encounter with Titan is a direct-entry arrival, 

phased to arrive at the satellite when Titan is receding 

from the spacecraft along the spacecraft’s approach 

asymptote (i.e. at a point in Titan’s orbit near the 

minimum spacecraft-Titan relative velocity). The 

arrival at Titan is indeed direct – there is no Saturn 

orbit insertion occurring prior to TOI (e.g. as will 

happen with Cassini). Titan’s north pole is targeted in 

order to place the spacecraft into a polar orbit. The 

orbiter’s closest approach to Saturn occurs 13 hours 

before Titan-arrival at a range of 1.2x106

 km 

(approximately equal to Titan’s orbital radius). 

The trajectory during approach is ballistic and the 

spacecraft depends upon chemical thrusters for flight 

path control. 

TARGET DESIGN 

The entry interface target consists of three 

parameters: inertial entry flight path angle (FPA), 

B-plane angle, and entry interface radius, where the 

interface entry radius is defined to equal 3575.0 km 

i.e. an altitude of 1000 km. The flight path angle is 

the angle subtended by the vehicle trajectory with the 

local horizontal at the EI, and it defines a corridor 

through the atmosphere (see Figure 4). The 

corridor’s width is constrained by upper and lower 

bounds determined by the physical limitations of the 

flight system (vehicle must be able to withstand 

aerodynamic, structural, and heat loads), and by the 

need to accumulate sufficient drag forces to slow the 

spacecraft (to avoid skip-out). 
38

For the orbiter, the entry interface target is derived 

from the characteristics of the vehicle aerodynamics 

and the post-insertion Titan orbit. For the probe, the 

entry interface target is derived from the desired 

landing site (latitude and longitude). 

The entry flight system at Titan is designed for an 

inertial entry velocity of 6.5 km/s and an entry FPA 

of -36.8°± ~1.0 (3σ) [4]. The error bars on the FPA 

are preliminary. The probe is designed to satisfy an 

entry FPA of -50°±5 (3σ) [4]. 

The atmosphere at Titan (mostly nitrogen and 

methane) has a surface density of 5 kg/m3

 (5x Earth), 

corresponding to a surface pressure of ~1.5 bar [5]. 

Subject to these conditions, an FPA of -50° subjects 

the probe to a maximum deceleration of ~10 g’s. An 

FPA of -36.8° subjects the orbiter to a maximum 

deceleration of ~4 g’s (assuming the baseline 

atmospheric density profile described in Reference 

6). 

The orbiter descends to an altitude of ~300 km before 

exiting the atmosphere (also a function of atmosphere 

density) [6]. Two hours after aerocapture a 

propulsive maneuver raises pericenter out of the 

atmosphere and circularizes the orbit. For an entry 

velocity of 6.5 m/s the pericenter-raise ∆v is 150 m/s 

(deterministic). The target and orbit characteristics 

(after circularization) are provided in Table 1. 

Table 1: Orbiter Entry Interface Target and 

Post-insertion Orbit Characteristics 

Entry time: Aug 20, 2016 23:56:18 UTC 

EI Target Orbit Characteristics 

Altitude 

(km) 

Latitude 

(deg) 

eFPA 

(deg) 

Entry 

Velocity 

(km/s) 

Altitude 

(km) 

Inclination 

(deg) 

Period 

(hours) 

1000 80N -36.8 6.5 1700 100 5 

NAVIGATION DATA 

Optical Data 

Optical data consists of digital images of Titan, other 

Saturnian satellites, and possibly Saturn, set in front 

of a stellar background. Background stars, combined 

with planetary ephemerides, establish the spacecraftTitan relative position by astrometry. And since 

radiometric data are unable to resolve a gravity 

signature from Titan until a spacecraft is nearly upon 

it, target imaging is important for approach 

navigation. 

The optical navigation campaign begins at E-75 days. 

(In practice, it’s more likely that navigation images 

would start as early as T-180 days.) Ground-based 

facilities will process transmitted pictures to extract 

the optical observables; then the data will be 

combined with radiometric measurements. 

Opnav transmissions will be constrained by the 

downlink data rate (600 bps). Probably more 

significant, however, are conflicts with competing 

spacecraft activities. Anticipating such demands 

during the last fortnight of cruise, a maximum rate of 

one opnav per every 2-4 hours was selected as a 

reasonable compromise [7]. 

At the beginning of the approach phase, one picture 

every other day is shuttered. (Titan’s orbital period is 

16 days.) The picture frequency increases to five per 

day, and increases again to approximately ten opnavs 

per day within 16 days of Titan. This yields 

approximately 200 images in the optical data set. 

The imaging system envisioned here follows a design 

similar to the Mars Reconnaissance Orbiter optical 

navigation camera. Relevant technical specifications 

of the advanced (MRO) camera are: aperture = 6 cm, 

focal length = 50 cm, field-of-view = 1.4° per side, 

detector = 1024x1024 CCD array, pixel resolution = 

50 μrad, mass = 2.7 kg, peak power = 4 W [8]. 

For comparison, the Cassini wide-angle camera has 

these specifications: aperture = 6 cm, focal length = 

20 cm, field-of-view = 3.5° per side, detector = 

1024x1024 CCD array, pixel resolution = 60 μrad, 

mass = 27 kg, peak power = 35 W [7]. The advanced 

(MRO) camera offers higher resolution yet weighs 

less and is more energy efficient. 

See Appendix 1 for optical data weights. 
39

Tracking Data 

The baseline navigation data set throughout the 

mission consists of two-way coherent X-band doppler 

and two-way coherent X-band range measurements. 

These data are augmented during approach with 

optical observations and ∆VLBI measurements.

∆VLBI data enhances the navigation solution relative 

to that achievable with doppler, range and optical data 

(although it is the optical data that dominates in a 

ranking of the relative importance of the four data 

types). In general ∆VLBI data, or specifically for this 

analysis Delta Differenced One-way Range (∆DOR), 

has limited effectiveness because of the spacecraft’s 

distant location in the solar system, but it can be used 

to great advantage in combination with other data 

types. 

Table 2: Doppler and Range Tracking Coverage 

Start End Coverage

E-75days E-41 3 tracks/week 

E-40 E-11 2 tracks/day 

E-10 Entry 3 tracks/day 

Data frequencies used in this analysis for doppler, 

range and ∆DOR are provided in Tables 2 and 3. Data 

weights are provided in Appendix 1. 

Table 3: DDOR Coverage 

Start End 

Observation 

Frequency

E-75days E-51 1 per week 

E-50 E-35 3.5 per week 

E-34 Entry 14 per week 

Navigation Performance 

Significant error sources in the navigation model are 

listed below. (Also see Appendix 1which lists all 

error sources and a priori uncertainties.) The 

combined effect of maneuver execution errors and 

orbit determination errors mapped to the atmospheric 

entry interface point is called the delivery accuracy. 

Ephemeris Determination 

Bounds on the Titan ephemeris errors in the year 

2016 are given in Table 4. The tabulation is given in 

a Saturn-centered RTN coordinate system, where R 

represents the radial direction, T the down-track 

direction and N the out-of-plane direction. 

Significant improvements to the Titan ephemeris 

between now and 2016 should occur. The suffixes 

BOC and EOC in Table 4 illustrate this improvement 

with respect to current ground-based observational 

knowledge (indicated by “GB”). “BOC” (beginning 

of Cassini) denotes Titan’s ephemeris knowledge 

shortly after Cassini’s arrival at Saturn. “EOC” (end 

of Cassini) represents the expected Titan ephemeris 

knowledge in 2008 after approximately 44 Cassini 

flyby encounters. The improvement is appreciable. 

Table 4: Titan, Saturn and Earth Ephemeris 

Uncertainties (3σ) 

Mapped to August 21, 2016 

Central Body

R 

(km)

DT 

(km) OOP RSS 

Earth 0.01 2.6 3.9 4.7 

Saturn GB* 95 405 135 437 

Saturn BOC 27 81 51 99 

Titan GB* 120 570 300 655 

Titan BOC 6 482 6 492** 

Titan EOC 1 30 1 104** 

* Representative of accuracies currently possible from groundbased observations. Not mapped to 2016. 

** Combined with Saturn BOC error. 

The BOC and EOC ephemeris accuracies in Table 4 

include secular degradation to 2016. 

Maneuver Determination 

Maneuvers scheduled during the approach phase are 

listed below. Table 5 defines the baseline targeting 

strategy adopted for this analysis (last maneuver at E24 hours), while Table 6 illustrates an alternative 

targeting strategy (last maneuver at E-6 hours). The 

important difference between the two strategies is the 

probe release time. For the strategy outlined in Table 

5, the probe is released at E-30 days whereas the 

alternative strategy releases the probe at E-5 days. 

(A late release time, i.e. E-5 days, is preferred in 

order to reduce probe instability growth after 

separation and achieve an accurate delivery.) 

The TCM3 magnitude, parameterized as a function of 

probe release time, is plotted in Figure 3. The 

remaining maneuvers in Tables 5 and 6 are small 

statistical maneuvers. The magnitudes of these 

clean-up maneuvers have not been analyzed, but the 

mean ∆v for TCM1 and TCM2 probably will not 

exceed 1 m/s each for either strategy. For the PR-30 

strategy, the combined total of TCM4 and TCM5 will 

be on the order of 5-10 m/s (at a confidence level of 

1σ). For the PR-5 strategy, the mean ∆v for TCM4 
40

and TCM5 will be much greater. 

Table 5: PR-30 Approach Phase TCMs 

TCM* Time** 

OD Data 

Cutoff** Description 

TCM-1 E -60 

days 

E - 65 

days 

Correct SEP cruise errors; 

penultimate probe targeting. 

TCM-2 E -31 

days 

E - 35 

days 

Final probe targeting to probeentry aim point. 

PR E -30 

days 

E – 30 

days Probe release. 

TCM-3 

(TDM) 

E -29 

days 

~E – 30 

days 

Deflect orbiter to entry 

interface point. 

TCM-4 E – 9 

days 

E – 11 

days 

Correct deflection maneuver 

errors; penultimate targeting. 

TCM-5 E – 24 

hrs 

E – 48 

hrs 

Final orbiter targeting to orbiterentry aim point. 

*Numbered starting at the beginning of the approach phase. 

 **With respect to orbiter entry (E) time. 

The 1σ execution errors assumed here are: fixed 

magnitude error of 1 mm/s, propor-tional magnitude 

error of 1% per axis, fixed pointing error of 1 mm/s 

per axis, and proportional pointing error of 2 

milliradians per axis for the deflection maneuver and 

1 milliradians per axis for the other TCMs. 

Orbit Determination 

The dominant orbit determination uncertainties 

consist of the reaction caused by the probe release 

mechanism, ephemeris errors, TCM execution 

uncertainties, and data errors. 

Table 6: PR-5 Approach Phase TCMs 

TCM Time 

OD Data 

Cutoff Description 

TCM-1 E –60 

days 

E - 65 

days 

Correct SEP cruise errors; 

penultimate probe targeting. 

TCM-2 E -6 

days E - 8 days Final probe targeting to probe-entry 

aim point. 

PR E -5 

days E – 5 days Probe release. 

TCM-3 

(TDM) 

E -4 

days 

~E – 5 

days 

Deflect orbiter to entry interface 

point. 

TCM-4 E -24 

hrs E – 48 hrs Correct deflection maneuver errors; 

orbiter targeting. 

TCM-5 E -6 hrs E – 15 hrs Final orbiter targeting to orbiterentry point. 

Results 

Delivery errors are a combination of orbit 

determination errors and maneuver execution errors. 

Results for both the orbiter bus and entry probe are 

presented here. Unless noted otherwise, all results 

assume EOC Titan ephemeris knowledge. Sensitivities examined were: delivery improve-ments due 

to optical navigation data and/or ∆DOR observations, 

and entry FPA uncertainties parameterized by probe 

release time and/or the orbiter over-flight delay. 

Orbiter Vehicle. Mission OD uncertainties in the 

Titan B-plane for a probe release time of E-30 days 

and an over-flight delay of 3.5 hours are shown in 

Figure 5. TDM is not shown in Figure 5; it lies off 

the left edge of the figure. But it’s the rapid 

reconstruction of TDM (i.e. falling FPA uncertainty 

shown in Figure 5) that enables aerocapture to be 

undertaken. 

Sensitivities are shown in Tables 7 - 10. Table 7 

illustrates the improvement in the overall delivery 

accuracy due to the incremental addition of advanced 

data types to the basic doppler and range data set. 

The results in Table 8 show the delivery sensitivity to 

probe release time. Table 9 shows the effect of 

varying the a priori Titan ephemeris. Table 10 lists 

the probe delivery accuracy. 

Figure 5: FPA v. Time-to-Go (3σ) 

Margin remains in the navigation sub-system. Table 

7 (and Figure 5) show the delivery requirement is 

satisfied by any tracking option, excluding the 

doppler and range only option. 

The probe release time influences the magnitude of 

TDM and the telecom link between the probe and 

bus. Greater orbiter eFPA uncertainties can be 

expected for large distances between the orbiter and 

probe (i.e. long over-flight delay times) and for late 

probe release times (i.e. releasing closer to Titan). 
41

This performance is illustrated in Table 8. 

Knowledge of Titan’s orbital position increases 

during Cassini’s tour (Table 4). This improvement is 

apparent in the flight path angle errors shown in 

Table 9. 

Table 7: Orbiter Delivery Accuracy (3σ) 

 Data Sensitivity

Probe release = E-30 days, Over-flight delay = 3.5 hours 

Doppler & 

Range 

Only 

Doppler 

Range & 

∆DOR 

Doppler 

Range 

& 

 Optical 

Doppler 

Range 

Optical

∆DOR 

Preliminary Reqm’t ±1.0 ±1.0 ±1.0 ±1.0 

Nominal Delivery

 

Data Cutoff at E-48 hours 

Semi-major axis (km) 68.7 39.1 28.1 23.3 

Semi-minor axis (km) 17.4 16.0 13.3 11.7 

Ellipse angle (deg) 75.2 66.0 64.9 53.7 

Entry time (s) 17.5 9.1 6.7 4.8 

B magnitude (km) 68.7 38.6 27.6 22.0 

Flight Path Angle (deg) ±1.7 ±1.0 ±0.7 ±0.6 

Alternate Delivery 

Data Cutoff at E-30 hours 

Semi-major axis (km) 37.0 30.0 21.9 20.9 

Semi-minor axis (km) 16.1 14.2 12.6 10.9 

Ellipse angle (deg) 65.3 55.0 57.4 48.0 

Entry time (s) 10.9 7.9 5.5 4.0 

B magnitude (km) 36.5 28.5 21.1 19.2 

Flight Path Angle (deg) ±0.9 ±0.7 ±0.5 ±0.5 

Parameter Update 

Data Cutoff at E-3 hours 

Semi-major axis (km) 15.8 7.2 13.1 7.1 

Semi-minor axis (km) 1.5 1.6 1.5 1.7 

Ellipse angle (deg) 56.2 57.2 56.3 57.9 

Entry time (s) 3.4 1.3 1.5 1.2 

B magnitude (km) 14.9 6.9 12.3 6.7 

Flight Path Angle (deg) ±0.4 ±0.2 ±0.3 ±0.2 

Table 8: Orbiter Delivery 

Probe Release Time v. Over-flight Delay 

Entry FPA Uncertainty (3σ) 

Probe Release 

E – 30 days 

Probe Release 

E – 5 days 

Over-flight Delay -> 3.5 hrs 2 hrs 2 hrs 1 hr

E – 48 hours (deg) ±0.6 ±0.6 ±1.1 ±1.0 

E – 30 hours (deg) ±0.5 ±0.4 ±0.9 ±0.8 

E – 3 hours (deg) ±0.2 ±0.2 ±0.6 ±0.6 

 

Table 9: Orbiter Delivery 

Probe Release Time v. Ephemeris Sensitivity 

Entry FPA Uncertainty (3σ) 

Probe Release = E - 30 days 

Over-flight Delay = 3.5 hours

Titan ephemeris ->

 circa 2008 

(EOC)

circa 2004

(BOC) 

circa 2000

(GB) 

E – 48 hours (deg) ±0.6 ±0.7 ±3.6 

E – 30 hours (deg) ±0.5 ±0.6 ±3.5 

E – 3 hours (deg) ±0.2 ±0.3 ±0.5 

The delivery dispersions of entry interface states 

resulting from monte carlo simulations of navigation 

errors are shown in Figure 6. The larger scatter (grey 

points) represents the B-plane uncertainty at the E-2 

day delivery, while the tighter scatter (black points) 

depicts the uncertainty at an E-3 hour delivery. 

The E-2 day scatter shown in Figure 6 maps into 

flight path angle uncertainty shown in Figure 7. 

Figure 7 plots FPA as a function of entry velocity for 

the E-2 day delivery. The scatter is not clustered 

around -36.8° because the navigation entry states are 

mapped to 1 minute before EI (radius = 3821 km) 

and not the EI (radius = 3575 km). 

Earlier Study. A navigation analysis undertaken in 

May 2002 generated preliminary results for the Titan 

aerocapture study group, as discussed in Reference 1. 

The results presented here represent more recent 

analyses. (In the earlier study, although the advanced 

camera was also investigated, it did not represent the 

baseline case. Ephemeris revisions have also 

occurred since then.) 

Probe capsule. The ultimate probe-targeting 

maneuver occurs thirty-one days before entry (in the 

baseline case). Subsequent to that maneuver the 

spacecraft spins up to prepare for probe release. At 

E-30 days the probe separates from the spacecraft 

bus. Contact with the probe is broken for 30 days 

while the probe follows a ballistic path to Titan. 

Upon reaching the entry interface point (radius equal 
42

to 3575 km) the probe begins broadcasting telemetry 

on UHF. The orbiter, at that time about 80,000 km 

behind the probe, begins relaying the probe’s data to 

Earth for the next 3.5 hours. Contact with the probe is 

lost at about the time the orbiter reaches its entry 

interface point. 

The requirement to deliver the probe to Titan can be 

satisfied, as shown in Table 10, for either an E-30 

day release or an E-5 day release. The latter probe 

release time delivers the probe to Titan with higher 

accuracy. The probe can separate as early as ~E-90 

days and still meet the delivery requirement. 

Table 10: Probe Delivery Accuracy (3σ) 

Probe Release Time and Camera Sensitivity

 Probe Release 

E – 30 days 

Probe Release 

E – 5 days 

Camera Design 

circa 

2008 

legacy 

camera 

circa 

2008 

legacy 

camera

Preliminary Requirement ±5.0 ±5.0 ±5.0 ±5.0 

Flight Path Angle (deg) ±1.9 ±1.9 ±0.8 ±0.9 

The probe model differs in minor ways from the bus 

model. The significant differences to note are the 

initial state uncertainties due to the probe-release 

mechanism, the absence of any trajectory correction 

maneuvers after separation, and the paucity of 

tracking data after separation. 

DISCUSSION 

Significant benefits accrue by assuming Titan’s 

position will be known to about the same level of 

certainty in 2021 as will exist at the time of Cassini’s 

last flyby of Titan. This is a reasonable assumption 

since Titan’s orbit has few perturbations and over a 

decade any error growth should be small (unlike the 

Galilean satellites of Jupiter). The sensitivity of the 

delivery to Titan’s ephemeris is shown in Table 9. 

Optical data is an important component in the 

navigation data suite, although its relative utility 

diminishes given precise and accurate Titan position 

knowledge. The EOC Titan uncertainty in Table 4 is 

nearly equal to the accuracy achievable with optical 

navigation (within a factor of ~2). Therefore opnavs 

contribute proportionally less to navigation early in 

the approach phase because errors other than Titan’s 

ephemeris dominate (e.g. error in the spacecraft 

state). See plot 2 in Figure 5. In addition to a 

measurable improvement to the delivery from 

incorporating the Cassini EOC-level Titan ephemeris, 

modest improvements in camera technology can also 

improve the delivery (the difference between plots 4 

and 5 in Figure 5). 

Optical navigation images of Titan, because of its 

atmosphere, have relatively large uncertainties, 

especially in the week before entry. This uncertainty 

is partially mitigated by incorporating navigation 

pictures of the small, icy moons of Saturn. 

∆DOR improves delivery accuracy (see Figure 5 and 

Table 7). This improvement in delivery accuracy 

comes about because the a priori Titan position 

uncertainty is less than the optical data noise (early in 

the approach phase). This means delivery knowledge 

is not dominated by ephemeris errors, but rather by 

the spacecraft’s state uncertainty. ∆DOR 

observations decrease spacecraft errors. Only within 

four days of Titan does optical data dominate the 

solution. 

∆DORs and optical data are orthogonally 

complementary and thus can combine to yield plot 4 

in Figure 5. This is the best delivery in the current 

study. Note that plot 5 represents a similar 

simulation, but with a less advanced camera. In the 

latter case optical data noise swamps ephemeris and 

spacecraft errors, resulting in less precise orbit 

determination. 
43

There is no advantage to using Ka-band tracking in 

place of X-band. There may be benefit to Ka-band 

∆DOR observations, but this parameterization has not 

yet been explored. 

 

Separation of the probe from the orbiter can occur as 

early as approximately 90 days before Titan and 

satisfy the probe’s entry requirements. However, at 

the opposite end, for a separation equal to E-5 days, 

several disadvantages become apparent. The 

maneuver schedule in the last week is compressed 

significantly with respect to a separation at E-30 days 

or earlier. The magnitude of the deflection maneuver 

(TDM) is large. This is offset with an accurate probe 

delivery to Titan. 

The data arc used for the E-30 day probe release 

consists of 30 optical navigation images and 32 

∆DOR observations. For the E-30 day release time, 

both the baseline case (MRO-like camera) and a less 

capable legacy imaging system provide equivalent 

probe deliveries, both well within the delivery 

requirement. 

Entry time uncertainty for the probe is 20 seconds 

(3σ) for the baseline case (i.e. for an E-30 day 

release, four navigation data types, EOC ephemeris). 

The relay link on the orbiter bus needs this margin 

for the telecom design. 

The Titan orbiter mission is boosted by SEP during 

early cruise, but during the latter half of the mission 

conventional thrusters control the spacecraft. As a 

rough estimate of propellant loading, for an E-30 day 

separation and a 3.5 hour over-flight delay, the 

spacecraft will need to carry propellant sufficient to 

perform about 300 m/s of velocity change, including 

both deterministic and statistical maneuvers (up to 

and including the post-insertion orbit circularization 

burn, but not including on-orbit maintenance 

propellant) [2]. 

Entry FPA results or expected results from other 

missions are summarized in the table below. (MER, 

Huygens, and Stardust have not yet arrived at Mars, 

Titan and Earth respectively at the time of this 

writing.) 

MPL and Huygens stand out in the short list above 

with high uncertainties. 

The MPL mission was characterized by unbalanced 

and mis-modeled thrusting activities. The level of 

thrusting required by the ACS system to maintain 

attitude significantly exceeded pre-launch 

expectations, and this mis-modeling contributed to 

the entry flight path angle uncertainty shown in Table 

11. 

Table 11: Delivery Accuracy Comparison (3σ) 

Orbiter and Probe v. Other Missions

 Entry 

FPA 

Error Delivery 

Time Reqm’t

Titan orbiter -36.8° ±0.6° E-2 d <±1.0> 

Mars Pathfinder -14.2° ±0.4° E-2 d ±1.0 

MPL -12.0° ±1.0° E-2 d ~±0.5 

MER -11.5° ±0.2° E-2 d ±0.25 

Stardust -8.2° ~±0.8° E-2 d ±0.8 

Titan probe -50.0° ±1.9° E-35 d ±5.0 

Galileo probe -8.6° ±0.6° E-140 d ±1.4 

Huygens probe -64.0° ±3.0° E-21 d ±3.4 

<-> indicates a proposed requirement 

Huygens will be released from Cassini 21 days 

before entry and, with an uncertainty of +/-3.0°, is 

not significantly different from the Titan orbiter 

probe delivery. The Huygens delivery is larger than 

the Titan probe because it does not have access to the 

Titan EOC ephemeris. Another reason is the tour redesign Cassini has undergone recently [9]. The 

release time is no longer as favorable for probe 

delivery, subjecting Huygens to additional 

perturbations. 

The MER delivery, on the other hand, is significantly 

smaller than MPL and the current delivery. The 

reasons include: well known Mars’ ephemerides, 

large gravity-well, lack of a TDM, and no force mismodeling. 

SUMMARY 

This study has baselined the use of optical 

observations and ∆DOR measurements in addition to 

doppler and range data for delivering a probe and 

orbiter to Titan, and has assumed a Cassini-era Titan 

ephemeris. The orbiter can satisfy its proposed entry 

FPA error requirement of ±1.0° (3σ) as early as 7 

days from Titan. By E-2 days that error has 

decreased by 170% (i.e. to ±0.6°). Nav-igation is 

sensitive to ∆DOR measurements, and the E-2 day 

delivery accuracy improves by 0.1° vis-à-vis only 

doppler, range and optical data, which is significant. 

An alternate delivery at E-30 hours reduces the E-2 

day FPA uncertainty by an additional 15%. 
44

Current camera technology (Cassini-era) is sufficient 

to meet the delivery requirement. More advanced 

cameras (MRO-era) will significantly improve the 

delivery and offer mass savings. 

The orbiter cannot meet preliminary entry 

requirements using a pre-Cassini ground-based a 

priori Titan ephemeris such as that described in 

Table 4. 

The pericenter-raise maneuver following orbit 

insertion is a significant issue that has not been 

addressed here. Maneuver autonomy may be 

required. 

This work represents a first-cut effort at determining 

concept feasibility. Many simplifying assumptions 

were made, especially with respect to the optical 

data, in order to accomplish this study in a timely 

manner. 

ACKNOWLEDGEMENTS 

This work was performed at the Jet Propulsion 

Laboratory, California Institute of Technology under 

contract from NASA. The work was funded through 

the In-space Propulsion program. 

The author would like to thank J.L. Hall, M. Noca, R. 

Bailey, T. Spilker, W.M. Owen, T. Goodson, P. 

Antreasian, D. Roth, P. Knocke, and L.A. 

Cangahuala for helpful comments and support. 

Many thanks also to R. Ionasescu for providing the 

simulated Cassini-era Titan ephemeris. Carl Sauer 

constructed the intial SEP trajectory between Earth 

and Saturn. 

REFERENCES 

1. Lockwood, M.K., “Titan Aerocapture Systems 

Analysis”, paper AIAA-2003-4799, Joint 

Propulsion Conference, Huntsville Alabama, 

July 21-23, 2003. 

2. Bailey, R.W., J.L. Hall, T. Spilker, “Titan 

Aerocapture Mission and Spacecraft Design 

Overview”, paper AIAA-2003-4800, Joint 

Propulsion Conference, Huntsville Alabama, 

July 21-23, 2003. 

3. Noca, M., JPL, personal communication, May 

2003. 

4. Lockwood, M.K., D.W. Way, Langley Research 

Center, personal communication, March 2002. 

5. Lorenz, R.D. et al, “The Seas of Titan”, EOS, 

Vol. 84, Number 14, 8 April 2003. 

6. Justus, C.G., “Aerocapture Systems Analysis 

Review”, viewgraph presentation, 29 Aug. 2002. 

7. Owen, W.M., JPL, personal communications, 

April 2002 - May 2003. 

8. Fraschetti, G., JPL, “MRO Optical Navigation 

Camera Preliminary Design Review”, viewgraph 

presentation, 8 May 2002. 

9. Strange, N., T. Goodson, Y. Hahn, “Cassini Tour 

Redesign for the Huygens Mission”, paper 

AIAA-2002-4720, AIAA Specialist Conference, 

Monterey Calif, 5-8 August 2002. 
45

 

 Appendix 1: A Priori Navigation Model Uncertainties 

Error Source A Priori 

Uncertainty 

(1σ) 

Correlati

on Time 

Comments 

Data 

doppler (mm/s) 0.1 - X-band 

range (m) 20 - Relatively high uncertainty in lieu of range 

biases 

∆DOR (nrad) 4 - 0.11 ns (X-band) 

optical (pixels) 0.25 - ~5.0 - Larger error corresponds to smaller range 

Estimated Parameters 

epoch state 

position (km) 1000 - 

velocity (km/s) 1 - 

Saturn ephemeris (km) (9, 27, 17) - R,AT,OOP,circa 2016 (5% DE405 error) 

Saturn mass (km3/s2) 0.0002 - circa 2016 (from simulated ephemeris) 

Saturn pole direction (mdeg) (0.12, 0.14) - R.A., Dec, circa 2016 (from simulated 

ephemeris) 

Titan ephemeris (km) (1, 10, 1) - R, AT, OOP, circa 2016 (from simulated 

ephemeris) 

camera pointing error (deg) (0.5, 0.5, 2) 0 R.A., Dec, Twist; estimated per observation 

non-gravitational accelerator 

(km/s2) 

1.0 x 1012 10 days Spherical covariance, estimated daily (1 day 

batches) 

solar pressure (%) 10 - Reflectivity coefficient 

ACS∆V (mm/s), 1 every 3 wks. (2, 2, 2) - (line-of-sight, lateral, normal) components 

TCMs (mm/s) spherical covariance 

TCM-1 4 - 

TCM-2 2 - 

proble release 5 - 

TCM-3 330 - 

TCM-4 5 - 

TCM-5 7 - 

3% (3σ) proportional error (per axis) 

3 mm/s (3σ) fixed error (per axis) 

Probe release at E – 30 days 

TCM3 at E – 29 days 

TCM4 at E – 9 days 

TCM5 at E – 1 day 

Earth pole direction (cm) 2→10 0 

UT1 (cm) 2→10 0 

(X and Y). Ramps to higher value during final 

week of data. 

(For UT1, ~10cm -> 0.26 ms.) 

ionosphere-day (cm) 55 0 

ionosphere-night (cm) 15 0 

troposphere (cm) 1 0 

S-band values 

Considered Parameters 

station locations (cm) 3 - 

quasar locations (nrad) 2 - for ∆DOR data 
46

Appendix 2: B-plane Description 

Planet or satellite approach trajectories are typically described in aiming plane coordinates referred to as “B–plane” 

coordinates (see Figure). The B-plane is a plane passing through the body center and perpendicular to the asymptote 

of the incoming trajectory (assuming two body conic motion). The "B-vector" is a vector in that plane, from the 

body center to the piercing-point of the trajectory asymptote. The B-vector specifies where the point of closest 

approach would be if the target body had no mass and did not deflect the flight path. Coordinates are defined by 

three orthogonal unit vectors, S, T, and R, with the system origin at the center of the target body. S is parallel to the 

spacecraft v∞ vector (approximately the velocity vector at the time of entry into the target body’s gravitational 

sphere of influence). T is arbitrary, but typically specified to lie in the ecliptic plane (the mean plane of the Earth’s 

orbit), or in the body equatorial plane. Finally, R completes an orthogonal triad with S and T. 

TARGET

PLANET

DISPERSION

ELLIPSE

TRAJECTORY

PLANE

AIMING PLANE 

("B-PLANE")

HYPERBOLIC 

PATH OF 

SPACECRAFT

B

INCOMING 

ASYMPTOTE 

DIRECTION

SMAA

SMIA

θ

T

R

DISPERSION ELLIPSE 

ORIENTATION

θ

S

T

R

Fig. Aiming Plane Coordinate System Definition 

Orbit determination errors can be characterized by a statistical dispersion ellipse in the aiming plane (B–plane) and a 

statistical uncertainty along the S (down-track) direction. In the Figure, SMIA and SMAA denote the semi–minor 

and semi–major axes of the dispersion ellipse (i.e. 50% of the distance across the ellipse along the respective 

coordinate). The angle θ is measured clockwise from T to SMAA. 
47 

ENGINEERING-LEVEL MODEL ATMOSPHERES FOR TITAN AND NEPTUNE 

C.G. Justus and Aleta Duvall 

Computer Sciences Corporation, Huntsville, AL 

D.L. Johnson

NASA Marshall Space Flight Center, Marshall Space Flight Center, AL 

ABSTRACT 

Global reference atmospheric models for Titan and 

Neptune have been newly developed for utilization in 

NASA's aerocapture systems analysis studies. Their 

applicability to the engineering design of guidance, 

navigation and control, and thermal protection systems 

is discussed. The legacy and key features of TitanGRAM and Neptune-GRAM are presented, with 

emphasis on the characterization of quasi-random 

atmospheric density perturbations. Sample Monte 

Carlo output for each model is presented. 

INTRODUCTION 

Engineering-level atmospheric models for Titan and 

Neptune have been newly developed for use in NASA’s 

systems analysis studies of aerocapture applications in 

missions to the outer planets. Analogous to highly 

successful Global Reference Atmospheric Models for 

Earth (GRAM)1

 and Mars (Mars-GRAM)2,3, the new 

models are called Titan-GRAM and Neptune-GRAM. 

Like GRAM and Mars-GRAM, an important feature of 

Titan-GRAM and Neptune-GRAM is their ability to 

simulate quasi-random perturbations for Monte Carlo 

analyses in developing guidance, navigation and control 

algorithms, and for thermal systems design. Figure 1 

compares density-height profiles for Earth, Mars, Titan, 

and Neptune. Relatively low scale heights (~10 km) 

make densities for Earth and Mars drop rather rapidly 

with altitude. Significantly higher scale height values 

for Titan and Neptune (~40 km) make these 

atmospheres considerably "thicker". Titan's large 

density scale height is due to its low gravity (~0.14 x 

Figure 1 - Typical density versus altitude on Earth, 

Mars, Titan, and Neptune. 
48 

Earth gravity), while that for Neptune is due to its low 

atmospheric mean molecular weight (~2.3 versus ~29 

for Earth). Vertical dotted lines in Figure 1 show 

density values and altitudes at which aerocapture or 

aerobraking maneuvers would occur on these planets. 

BASIS FOR THE ATMOSPHERIC MODELS 

In GRAM and Mars-GRAM, input values for date, 

time, latitude, longitude, etc. are used to calculate 

planetary position and solar position, so that effects of 

latitude variation, and seasonal and time-of-day 

variations can be computed explicitly. 

A simplified approach is adopted in Titan-GRAM and 

Neptune-GRAM whereby these effects (as well as 

effects of relatively large measurement uncertainties for 

these planets) are represented within a prescribed 

envelope of minimum-average-maximum density 

versus altitude. Figure 2(a) shows this envelope for 

Titan, for which engineering atmospheric profiles of 

Yelle et al.4

 are used. For Neptune, data from 

Cruikshank5

 were employed to generate a comparable 

envelope, shown in Figure 2(b). 

A single model input parameter (Fminmax) allows the 

user of Titan-GRAM or Neptune-GRAM to select

 (a) 

 (b) 

Figure 2 – Minimum, average, and maximum density versus altitude for 

(a) Titan4

 and (b) Neptune (developed from data in Cruikshank5). 
49 

where within the min-max envelope a particular 

simulation will fall. Fminmax = -1, 0, or 1 selects 

minimum, average, or maximum conditions, 

respectively, with intermediate values determined by 

interpolation (i.e. Fminmax between 0 and 1 produces 

values between average and maximum). Effects such 

as variation with latitude along a given trajectory path 

can be computed by user-selected representations of 

variation of Fminmax with latitude. 

Yelle et al. assumed methane mole fractions of 5%, 3%, 

and 1% for minimum, average, and maximum density 

Titan atmospheres, respectively. Aerothermal analyses 

(other papers in this session) showed very strong 

radiational heating by CN radicals produced from 

methane and nitrogen reactions in the high-temperature 

aerocapture entry shock layer. Because of the 

importance of methane concentration in this process, an 

option was added in Titan-GRAM to allow the user to 

select any amount of methane (up to 5%), while 

retaining the original Yelle et al. profiles of mean 

molecular weight versus height and density versus 

height. Atmospheric density perturbations are 

computed by methods similar to those used in GRAM 

and Mars-GRAM. Perturbation magnitudes in TitanGRAM and Neptune-GRAM are estimated from a 

methodology similar to that of Strobel and Sicardy6

, 

based on expected wave saturation effects. 

In particular, perturbation magnitudes are modeled 

from an approximate fit to the wave saturation 

condition relation 

ρ′max /ρ0 = [L/(2 π H)] ( 1 + [L/(2 π H)]2

)

1/2 

( dT0/dz + g/Cp ) / ( g/R ) 

where H is pressure scale height ( R T0 / g ), R is gas 

constant, g is acceleration of gravity, T0 is background 

mean temperature, and Cp is specific heat at constant 

pressure. Model perturbation magnitudes are illustrated 

in Figures 3(a) and 3(b). 

Titan Density Perturbation Magnitudes

0

5

10

15

20

25

30

35

40

0 100 200 300 400 500

Height, km

 Perturbation Magnitude, % of Mean

Equation, Yelle Avg., L/H=2

Stellar Occultation Estimate

Titan-GRAM Model

Figure 3(a) – Perturbation magnitudes in Titan-GRAM. Model values approximate equation (1), with ratio 

L/H=2 and dT0/dz from respective average atmospheres. Textured line is estimated value from Figure 1(a) of 

Strobel and Sicardy6

. 
50 

Figure 3(b) – Perturbation magnitudes in Neptune-GRAM. Model values approximate equation (1), with 

ratio L/H=2 and dT0/dz from respective average atmospheres. 

SAMPLE MODEL RESULTS 

Sample Monte Carlo outputs from Titan-GRAM and 

Neptune-GRAM are shown in Figure 4. 

CONCLUSIONS

Titan-GRAM and Neptune-GRAM are engineeringlevel atmospheric model for Titan and Neptune, 

suitable for a wide range of mission design, systems 

analysis, and operations tasks. For orbiter missions, 

Titan-GRAM and Neptune-GRAM applications include 

analysis for aerocapture or aerobraking operations, 

analysis of station-keeping issues for science orbits, 

analysis of orbital lifetimes for end-of-mission 

planetary protection orbits, and atmospheric entry 

issues for accidental break-up and burn-up scenarios. 

For Titan lander missions, Titan-GRAM applications 

include analysis for entry, descent and landing (EDL), 

and guidance, navigation and control analysis for 

precision landing. Using Titan-GRAM or NeptuneGRAM perturbation models in Monte Carlo mode 

make them especially suited for design and testing of 

guidance, navigation, and control algorithms and for 

heat loads analysis of thermal protection systems. 

ACKNOWLEDGEMENTS 

The authors gratefully acknowledge support from the 

NASA/Marshall Space Flight Center In-Space 

Propulsion Program. Particular thanks go to Bonnie 

James (MSFC), Manager of the Aerocapture 

Technology Development Project, to Michelle M. 

Munk (LaRC/MSFC), Lead Systems Engineer for 

Aerocapture, and to Ann Trausch and Melody 

Herrmann (MSFC), team leads and Mary Kae 

Lockwood (LaRC), technical lead for the 

Titan/Neptune Systems Analysis study. Model user 

feedback and suggestions from the following 

individuals are also greatly appreciated: Dick Powell, 

Brett Starr, and David Way (NASA LaRC), and Claude 

Graves and Jim Masciarelli (NASA JSC).

 

Neptune Density Perturbation Magnitudes

0

5

10

15

20

25

30

35

-100 100 300 500 700 900

Height, km

Max Perturbation Magnitude, % of 

Mean

Equation & Neptune Avg. L/H=2 Model
51 

(a) 

(b) 

Figure 4 - Sample Monte Carlo perturbation profiles from (a) Titan-GRAM and (b) Neptune-GRAM for 

Fminmax = -1, 0 and 1. Density values are expressed as percentage deviation from average density (Fminmax 

= 0) profile value.
52 

REFERENCES 

1

Justus, C.G., Johnson, D.L., and James, B.F., 

"New Global Reference Atmospheric Model 

(GRAM-99) and Future Plans", COSPAR 33rd 

General Assembly, Warsaw, Poland, Invited Paper 

C4.2-0003, July 2000. 

2

Justus, C.G., and Johnson, D.L., Mars Global 

Reference Atmospheric Model 2001 Version (MarsGRAM 2001) Users Guide, NASA/TM-2001-210961, 

April 2001. 

3

Justus, C.G., James, B.F., Bougher, S.W., 

Bridger, A.F.C., Haberle, R.M., Murphy, J.R., and 

Engel, S., "Mars-GRAM 2000: A Mars Atmospheric 

Model for Engineering Applications", Advances in 

Space Science, Vol. 29, No. 2, 2002, pp. 193202. 

4

Yelle, R.V., Strobel, D.F., Lellouch, E., and 

Gautier, D., "Engineering Models for Titan's 

Atmosphere", Huygens Science, Payload and 

Mission, ESA SP-1177, Aug 1997. 

5

Cruikshank, D.P., Ed., Neptune and Triton, 

University of Arizona Press, Tucson, 1995. 

6 Strobel, D.F., and Sicardy, B., "Gravity Waves 

and Wind Shear Models", Huygens Science, Payload 

and Mission, ESA SP-1177, Aug. 1997. 
53 

GUIDANCE ALGORITHMS FOR AEROCAPTURE AT TITAN 

James P. Masciarelli

NASA Johnson Space Center 

Houston, TX 

Eric M. Queen 

NASA Langley Research Center 

Hampton, VA 

ABSTRACT 

An evaluation of two different guidance algorithms for 

aerocapture at Titan is presented for a reference 

mission. The evaluation includes guidance algorithm 

comparison to the theoretical optimum performance, as 

well as guidance response to variation in entry flight 

path angle, atmosphere density, and aerodynamic 

parameters. Monte Carlo simulation results are also 

presented. The results show that both algorithms, 

originally developed for aerocapture at Earth and Mars, 

perform well for aerocapture at Titan. 

INTRODUCTION 

NASA is investigating the use of aerocapture for orbit 

insertion at Titan, one of Saturn’s moons.1

 An 

autonomous guidance algorithm will be required to 

perform this maneuver and achieve the desired orbit. 

The Hybrid Predictor-corrector Aerocapture Scheme 

(HYPAS), and the Terminal Point Controller (TPC) are 

two algorithms that have been developed for 

aerocapture guidance. Both algorithms have previously 

been tested extensively in simulations for aerocapture at 

Earth and Mars, but no investigation has been done on 

their performance at Titan. This paper explores the 

performance of these two algorithms for aerocapture at 

Titan to determine the feasibility of aerocapture for a 

reference mission. 

The proposed Titan Explorer was used as the reference 

mission to assess guidance performance for aerocapture 

at Titan.2

 In the reference mission, the vehicle enters 

the Titan atmosphere at an altitude of 1000 km and an 

inertial velocity between 6.5 and 10.0 km/s, depending 

on where in its orbit Titan is encountered. The vehicle 

utilizes a blunt body aeroshell, with lift-to-drag ratio 

(L/D) of 0.25, and ballistic coefficient (m/CDA) of 

90 kg/m2

. Lift modulation by bank angle control is 

used to manage the aerocapture trajectory. The target 

orbit is circular, with an altitude of 1700 km, and 

inclination of 101.6 deg. 

The paper begins with an overview of each guidance 

algorithm, and points to other references that describe 

the algorithms in more detail. The process used to 

analyze the guidance algorithms for the reference Titan 

aerocapture mission is then explained. Finally, the 

performance of the two algorithms is presented. A 

companion paper presents more detailed aerocapture 

performance results for the reference mission.3

OVERVIEW OF GUIDANCE ALGORITHMS 

HYPAS 

The HYPAS algorithm guides a lifting vehicle through 

the atmosphere to a desired exit orbit apoapsis and 

inclination or plane using bank angle control. The 

guidance uses an analytically derived control algorithm 

based on deceleration due to drag and altitude rate error 

feedback. Inputs to the guidance algorithm are the 

current vehicle position, velocity, sensed acceleration, 

and vehicle attitude. The algorithm outputs a 

commanded bank angle and the direction to bank from 

the current attitude. The guidance algorithm is 

adaptable to a wide range of initial state vectors, vehicle 

lift-to-drag ratios and ballistic coefficients, planetary 

atmospheres, and target orbits by changing a set of 

initialization constants. Furthermore, by tuning these 

constants, other trajectory constraints can be controlled 

such as maximum dynamic pressure, deceleration, heat 

rate, and the amount of the theoretical corridor 

captured. 

A significant feature of the HYPAS algorithm is that no 

reference trajectories are computed prior to flight; all 

reference values are computed and updated during 

flight. This analytic, “on-the-fly” approach leads to 

efficient code, minimal data storage requirements, and 

minimal preflight effort. The current implementation of 

the guidance algorithm uses 313 source lines of Fortran 

code, including code required to process navigation 
54 

data. There are 37 initialization constants, most of 

which are functions of the vehicle’s lift-to-drag ratio 

and ballistic coefficient, the planet’s atmosphere and 

gravity, and the target orbit conditions. The nonnumerical, non-iterative scheme ensures fast and 

consistent execution times. 

The HYPAS guidance algorithm consists of two phases. 

In the first phase, or capture phase, bank angle 

commands are generated to stabilize the trajectory and 

drive the vehicle toward equilibrium glide conditions, 

where lift, gravity, and centripetal forces are balanced. 

When the vehicle has decelerated to a specified 

velocity, the second phase, or exit phase, begins. In the 

exit phase, the velocity vector at atmospheric exit 

altitude is analytically predicted each guidance 

computation cycle. The bank angle command is then 

adjusted so that the velocity achieved at exit altitude 

will produce an orbit with the target apoapsis. This 

two-phase approach allows separate tuning of 

initialization constants to maximize robustness during 

capture and maximize performance during exit. 

An inclination (or wedge angle if targeting a specific 

plane) dead-band that is a function of inertial velocity is 

used to target the desired orbit inclination (or plane). 

Whenever the inclination error exceeds this dead-band, 

a bank reversal is commanded. The direction to bank is 

selected through a series of tests that examine current 

velocity, angular distance to roll, difference between 

desired and measured altitude rate, and difference 

between desired and measured drag. 

The HYPAS guidance algorithm has undergone 

extensive laboratory development and testing. The 

original version of the algorithm was developed for the 

Aeroassist Flight Experiment (AFE) program4

, and its 

derivation was published in Reference 5. During the 

AFE program, the algorithm was tested, compared, and 

evaluated against other guidance algorithms in three 

and six degree-of-freedom computer-based 

simulations.6

 The HYPAS guidance algorithm was 

selected for the space flight test, and development of 

the flight code was on schedule until the AFE program 

was cancelled. 

The HYPAS algorithm has been used in numerous 

human and robotic exploration mission studies 

performed at NASA over the last several years. These 

studies involved developing nominal and dispersed 

trajectory simulation results for aerocapture at Earth 

and Mars, for a wide range of vehicle L/D, ballistic 

coefficients, entry conditions, and target orbits. This 

work has provided the opportunity to gain a deep 

understanding of how the guidance algorithm performs 

in a variety of situations. Modifications have been 

made as necessary to improve performance and 

robustness. These modifications include maintaining 

the equilibrium glide drag reference into the exit phase, 

correction for measured L/D (from sensed acceleration 

vector) in the commanded bank equation, smoothing 

the transition between the capture and exit phases, and 

calculation of bank commands before atmosphere entry 

based on the estimated position in the entry corridor. 

TPC 

The TPC algorithm also uses bank control to guide a 

lifting vehicle to a desired apoapsis and inclination or 

plane. TPC is based on a calculus of variations 

approach and is analogous to the Apollo Earth-entry 

guidance, but with different boundary conditions. TPC 

is a feedback guidance that uses sensitivities of the exit 

condition to changes in the state and control to 

determine the control at any point along the trajectory. 

The sensitivities are generated from a reference 

trajectory that is determined off-line prior to flight. The 

guidance does not attempt to follow the reference 

trajectory or any other particular trajectory. The 

reference trajectory is simply used to generate the 

sensitivity coefficients. This reference trajectory is run 

open-loop with a reduced lift coefficient. Reducing the 

lift coefficient gives the same effect in apoapsis as a 

nonvertical bank angle without the out-of-plane effects. 

Any valid trajectory can be used as a reference, but 

usually one with a constant bank angle that meets the 

target condition is chosen. 

The TPC algorithm has been studied extensively for 

several proposed aerocapture missions. It was 

originally developed for the Mars Surveyor Program 

2001 Orbiter. The French space agency (CNES) ranked 

it best among several algorithms proposed for the 

“Premier” sample return mission before aerocapture 

was dropped from the mission. The analogous lander 

guidance was chosen for the Apollo Earth-entry 

guidance because it needed very little onboard 

computer resources, it was very accurate and it 

degraded gracefully under extreme conditions. 

TPC has an in-plane component, which targets the 

velocity increment (∆V) required to achieve a desired 

orbit after the atmospheric pass is complete, and an outof-plane component, which targets inclination or wedge 

angle. The out-of-plane logic is structurally similar to 

the HYPAS out-of-plane logic and relies on a deadband 

to trigger bank angle reversals. 

Derivation of the gains used by the TPC algorithm is 

presented in Reference 7. A reference trajectory is 

generated off-line. The adjoint equations to the 
55 

equations of motion are then integrated from the 

atmospheric exit backwards to the atmospheric entry. 

The adjoint variables (sensitivity coefficients) are used 

to create gains on drag acceleration, altitude-rate, and 

velocity. A data point reduction algorithm is then 

applied to the gains to reduce the size of the onboard 

data required. Note that this process is completely 

automated after the reference trajectory is chosen. 

On board, the guidance basically does a table lookup 

and multiplies the stored gains by the appropriate states. 

This determines a bank angle that will drive the final 

∆V to be near the reference ∆V. The entire trajectory is 

flown the same way and is treated as a single 

continuous passage. For cases with an extremely low 

apoapsis, an override command is applied to help raise 

the apoapsis. 

Density estimation of the atmosphere is also performed 

on board. The TPC density estimation uses a least 

squares fit of the most recent acceleration 

measurements in a manner which is theoretically very 

similar to that used by HYPAS, though the two 

implementations are unique. 

RECENT TESTING OF THE HYPAS AND TPC 

ALGORITHMS 

During 1998, the HYPAS and TPC algorithms were 

investigated for use on the Mars Surveyor Program 

2001 mission. The algorithms were tested in computerbased simulation environments at JSC and LaRC and 

found to perform well under nominal and dispersed 

conditions. The results from that work are published in 

Reference 7 and Reference 8. The algorithms 

underwent extensive analysis and testing for the Mars 

Sample Return Orbiter and the Mars Premier Mission 

studies that were jointly conducted with the French 

Space Agency, CNES. Again, the algorithms were 

found to perform very well in computer based 

simulations. The results from that work are published 

in References 9, 10, and 11. 

Although the performance of these algorithms has been 

studied extensively for Earth and Mars, there has been 

no previous analysis of their capabilities and 

performance for aerocapture at other destinations. The 

remainder of this paper describes a recent assessment 

for aerocapture at Titan. 

PERFORMANCE FOR TITAN AEROCAPTURE 

AEROCAPTURE ANALYSIS PROCESS 

The aerocapture guidance analysis process involves 

several steps. First, the theoretical entry flight path 

angle corridor is determined. This is used to establish 

the nominal target entry conditions and sets the bounds 

on the entry delivery errors for the given vehicle’s 

capabilities. Next, the theoretical optimum aerocapture 

performance is determined. This provides a benchmark 

for comparison with the performance obtained with the 

guidance algorithms. A nominal guided aerocapture 

trajectory is then developed. This involves an iterative 

process of adjusting guidance algorithm parameters and 

testing against variations in entry flight path angle, 

aerodynamic coefficients, and atmosphere density to 

get the best performance from the algorithm. Finally, 

Monte Carlo simulations are performed to evaluate the 

guidance performance under expected flight conditions. 

A computer program that simulates atmospheric flight 

about a central body through three-dimensional space, 

plus rotation about the vehicle’s velocity vector (bank 

angle) was used for this analysis. The central body, 

Titan, was assumed to have a spherical shape of radius 

2575 km, and an inverse square gravity field, with 

gravitational parameter of 9142 km3

/s2. Aerodynamic 

coefficients were assumed to be constant. The 

simulation uses an engineering model of the Titan 

atmosphere, which provides the capability to vary the 

density between minimum and maximum density 

profiles, as well as superimpose random density 

perturbations to simulate flight through a realistic 

atmosphere.12 Further details of the trajectory 

simulation used for this analysis are explained in 

Reference 3. 

ENTRY CORRIDOR 

The first step in developing a guided aerocapture 

trajectory is to determine the theoretical entry flight 

path angle corridor. The steep side of the corridor is 

defined to be the flight path angle at which the vehicle 

just reaches the apoapsis target with full lift up for the 

duration of the atmospheric flight. The shallow side of 

the corridor is defined to be the flight path angle at 

which the vehicle just stays below the target apoapsis 

with full lift down for the entire duration of the 

trajectory. The theoretical entry flight path angle 

corridor for the reference mission was determined using 

Table 1. Entry Corridors for Titan Aerocapture 

Atmosphere 

Profile 

Steep 

Side 

(deg) 

Shallow 

Side 

(deg) 

Width 

(deg) 

Middle 

(deg) 

Nominal -37.62 -34.15 3.46 -35.88 

Minimum -38.37 -35.28 3.09 -36.82 

Maximum -36.78 -32.91 3.87 -34.85 

Combined -36.78 -35.28 1.50 -36.03 
56 

the nominal, minimum, and maximum density 

atmosphere profiles and is shown in Table 1. Taking 

into account the total uncertainty in the atmosphere 

profiles, the entry flight path angle corridor width is 

1.5 deg. This is sufficient for the expected entry 

delivery errors.13 The nominal entry flight path angle 

was chosen to be the middle of the combined corridors, 

which is at -36.03 deg. 

THEORETICAL OPTIMUM PERFORMANCE 

With the theoretical entry corridors defined, the 

theoretical optimum performance can be determined. 

The theoretical optimum guidance exactly achieves the 

target apoapsis while maximizing the periapsis altitude 

(thus minimizing the post-aerocapture ∆V required) for 

all flight path angles within the theoretical corridor. 

This is achieved by entering the atmosphere with full 

lift vector up, and then switching to full lift vector 

down at the correct instant to reach the target apoapsis 

altitude. The point at which the switch from lift vector 

up to lift vector down occurs is a function of the entry 

flight path angle. For entry at the shallow side of the 

corridor, the switch to lift vector down occurs 

immediately at entry. For entry at the steep side of the 

corridor, the switch to lift vector down is not 

performed. The theoretical optimum guidance 

performance is easy to determine, but it is not practical 

to achieve in real flight. The usefulness of knowing the 

optimum performance is to evaluate how well practical 

guidance algorithms perform compared to the 

theoretical optimum. 

The theoretical optimum performance was determined 

for the reference mission and the nominal atmosphere 

profile for entry flight path angles across the theoretical 

corridor. The theoretical minimum ∆V required to 

circularize the orbit after aerocapture ranges from 

224 m/s at the steep side of the entry corridor to 

149 m/s at the shallow side of the corridor. For entry at 

the middle of the corridor (-36 deg), the periapsis 

altitude achieved is 290 km, which requires an ideal ∆V 

of 152 m/s after aerocapture to circularize the orbit at 

1700 km altitude. 

Once the theoretical corridor and optimum guidance 

performance were defined, the performance of the 

HYPAS and TPC algorithms were evaluated. 

HYPAS PERFORMANCE 

The HYPAS guidance algorithm parameters were 

adjusted for the reference aerocapture mission, and a 

nominal aerocapture trajectory developed. Figure 1 

shows the results from the nominal trajectory. 

Trajectory simulations with the HYPAS guidance 

algorithm were run with the entry flight path angle 

varied across the theoretical corridor for the nominal 

atmosphere profile. The results can be seen in Figure 2, 

which shows apoapsis altitude achieved and post0 2 4 6 8

−200

−100

0

100

200

Velocity, km/s

Bank Angle, deg

0 2 4 6 8

200

400

600

800

1000

1200

1400

Velocity, km/s

Altitude, km

0 2 4 6 8

100

101

102

103

104

105

Velocity, km/s

Orbit Inclination, deg

0 2 4 6 8

0

0.5

1

1.5

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Velocity, km/s

Sensed Acceleration, g

0 2 4 6 8

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0.5

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2.5

Velocity, km/s

Dynamic Pressure, kPa

0 2 4 6 8

0

10

20

30

40

Velocity, km/s

Convective Heat Rate, W/cm2

Figure 1. Nominal Guided Aerocapture Trajectory 
57 

aerocapture ∆V required versus entry flight path angle 

for the HYPAS and theoretical optimal guidance. The 

guided entry corridor was found to extend from -37.6 to 

-34.2 deg, which covers 98 percent of the theoretical 

corridor. The post-aerocapture ∆V required with the 

HYPAS guidance at the middle of the corridor is 

180 m/s, which is only 18 percent higher than the 

theoretical minimum. 

Trajectory simulations with the HYPAS guidance 

algorithm were also run with the entry flight path angle 

varied across the theoretical corridor for the minimum 

and maximum atmosphere profiles to determine the 

sensitivity to variation in the atmosphere density. 

Initially, it was found that the algorithm’s accuracy 

degraded significantly with the minimum and 

maximum density profiles. The reason for this is that 

the HYPAS algorithm assumes that the atmosphere 

density varies exponentially with altitude to predict exit 

conditions. The density scale height used in this model 

is a constant that is input prior to flight. This approach 

works well for missions where the atmosphere scale 

height does not vary significantly. However, the scale 

height of the Titan atmosphere does vary significantly 

between the minimum and maximum expected density 

profiles. The scale height for the nominal profile is 

approximately 50 km; for the minimum and maximum 

profiles, the scale height is approximately 42.5 and 

53 km, respectively. This is a variation of –15 to +6 

percent from the nominal and was found to be the 

source of the error in apoapsis altitude achieved. 

The variation in atmosphere scale height can be 

accounted for using on-board estimation of the 

atmosphere scale height from navigated altitude and 

sensed acceleration.14 A least squares filter was added 

to the HYPAS logic to do this. The results with and 

without the scale height estimation are shown in 

Figure 3. As can be seen, with the scale height 

estimation incorporated, the HYPAS guidance is 

insensitive to variation in the atmosphere density 

profile. 

An assessment of the HYPAS algorithm performance 

when subject to uncertainties in aerodynamic 

coefficients was made. Trajectory simulations were run 

with the entry flight path angle varied across the 

theoretical corridor for variations of ±10 percent in the 

lift and drag coefficients. The results are shown in 

Figure 4. As can be seen, the algorithm is insensitive to 

the variation in aerodynamic coefficients, which is due 

to the automatic adjustment in the bank angle command 

the guidance makes using the sensed acceleration. 

Monte Carlo simulations, which combine random 

variations in entry position and velocity vectors, 

aerodynamic coefficients, vehicle mass, mean 

atmosphere density profile, plus random atmosphere 

density perturbations, were performed to complete the 

assessment of the HYPAS algorithm’s performance for 

the reference Titan aerocapture mission. Figure 5 

shows a sample of the results for a 2000 case Monte 

Carlo with a nominal entry velocity of 6.5 km/s. 

Several other sets of Monte Carlo simulations were run 

with varying sets of assumptions. The results of those 

simulations can be found in Reference 3. The guidance 

algorithm was found to perform very well for the 

reference Titan aerocapture mission. 

TPC PERFORMANCE 

A nominal reference trajectory for the titan aerocapture 

was run and used to generate gains for the TPC 

guidance algorithm. The bank angle profile of the 

nominal trajectory is shown in Figure 6. The reference 

trajectory used a constant bank angle of about 62 

degrees.

−38 −37 −36 −35 −34

1400

1500

1600

1700

1800

1900

2000

Entry Flight Path Angle, deg

Apoapsis Altitude, km

Guided

Optimum

−38 −37 −36 −35 −34

150

200

250

300

350

Entry Flight Path Angle, deg

In−Plane ∆ V, m/s

Guided

Optimum

Figure 2. HYPAS Performance Compared to Theoretical Optimum 
58 

−39 −38 −37 −36 −35 −34 −33

1400

1500

1600

1700

1800

1900

2000

Entry Flight Path Angle, deg

Apoapsis Altitude, km

No Estimation of Scale Height

Nominal

Max

Min

−39 −38 −37 −36 −35 −34 −33

1400

1500

1600

1700

1800

1900

2000

Entry Flight Path Angle, deg

Apoapsis Altitude, km

With Estimation of Scale Height

Nominal

Max

Min

Figure 3. Effect of On-Board Estimation of Scale Height 

−38 −37 −36 −35 −34

1400

1500

1600

1700

1800

1900

2000

Entry Flight Path Angle, deg

Apoapsis Altitude, km

−38 −37 −36 −35 −34

150

200

250

300

350

Entry Flight Path Angle, deg

In−Plane ∆ V, m/s

Figure 4. HYPAS Sensitivity to Variation in Aerodynamic Coefficients 

−200 −100 0 100 200 300

1300

1400

1500

1600

1700

1800

1900

2000

2100

2200

Apoapsis Altitude, km

Periapsis Altitude, km

101 101.2 101.4 101.6 101.8 102

150

160

170

180

190

200

210

220

230

240

250

Orbit Inclination, deg

Post−Aerocapture ∆V, m/s

Figure 5. Monte Carlo Results Using HYPAS Guidance 
59 

For these initial test runs, a simple “funnel” deadband is 

used. The deadband is linear with velocity until a 

predefined velocity trigger (in this case 3000 m/s) at 

which point, the deadband switches to a constant width. 

The inclination history is also shown in Figure 6. 

Monte Carlo simulations, as for the HYPAS algorithm, 

were performed to complete the assessment of the TPC 

algorithm. Figure 7 shows the resulting periapsis and 

apoapsis when the vehicle left the atmosphere for 2000 

cases. None of these cases failed to capture, either high 

(skip out) or low (crash). A strong trend is obvious that 

the cases with high periapsis also have high apoapsis. 

The primary factor affecting the high cases seems to be 

initial flight path angle. All of the cases that ended with 

high apoapsis had a shallow entry flight path angle, 

though many cases with a shallow entry angle ended 

with apoapsis near the nominal. 

Figure 7 also shows inclination and ∆V to circularize 

the orbit at 1700 km. The deadband on inclination has 

performed well. All cases are within 0.5 deg of the 

target inclination of 101.6 deg. Most of the cases have 

a ∆V near 180 m/s, though a significant minority are 

clustered near 190 m/s. 99 percent of the cases have a 

final ∆V less than 201 m/s. 

SUMMARY AND CONCLUSIONS 

An evaluation of two different guidance algorithms, the 

HYPAS and the TPC, has been completed as part of an 

assessment of the feasibility of aerocapture at Titan. 

This evaluation included determination of the 

theoretical entry corridor and comparison of the 

0 500 1000 1500 2000 2500 3000

−200

−150

−100

−50

0

50

100

150

200

Time (sec)

Bank Angle (deg)

Bank Angle 

Bank Command

Titan Aerocapture, TPC Nominal Aeropass

0 2000 4000 6000 8000 10000 12000

99

99.5

100

100.5

101

101.5

102

102.5

103

103.5

104

Relative Velocity (m/s)

Inclination (deg)

Titan Aerocapture, TPC Nominal Aeropass

Figure 6. Nominal Bank and Inclination Profile Using TPC Guidance 

−50 0 50 100 150

1000

1500

2000

2500

Apoapsis vs. Periapsis

Periapsis Altitude (km)

Apoapsis Altitude (km)

TPC for Titan Aerocapture

101 101.2 101.4 101.6 101.8 102

165

170

175

180

185

190

195

200

205

99% of Delta V < 201.6627

Inclination (deg)

∆ V to circularize at 1700 km

Inclination vs. ∆ V TPC for Titan Aerocapture

Figure 7. Monte Carlo Results Using TPC Guidance 
60 

guidance algorithms’ performance to the theoretical 

optimum. Guidance responses to off-nominal 

conditions, including Monte Carlo simulations, were 

also investigated. 

The HYPAS and TPC algorithms have been previously 

studied for aerocapture at Earth and Mars, and they 

were used at Titan without any significant modification. 

The only difference is that on-board estimation of the 

atmosphere density scale height was included to 

improve guidance accuracy. This is because there is a 

large variation in current estimates of the scale height 

of Titan’s atmosphere. 

Both aerocapture guidance algorithms were shown to 

perform well for the proposed Titan Explorer mission. 

No major technical issues are expected in implementing 

a guidance algorithm for this reference mission. 

Therefore the Titan aerocapture mission is feasible 

from a guidance algorithm perspective. 

REFERENCES 

1. M.K. Lockwood, “Titan Aerocapture Systems 

Analysis,” AIAA-2003-4799, 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Huntsville, AL, July 2003. 

2. R. Bailey, J. Hall, T. Spilker, “Titan Aerocapture 

Mission and Spacecraft Design Overview,” AIAA2003-4800, AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference, Huntsville, AL, July 2003. 

3. D. Way, R.W. Powell, J. Masciarelli, B. Starr, K.T. 

Edquist, “Aerocapture Simulation and Performance 

for the Titan Explorer Mission,” AIAA-2003-4951, 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Huntsville, AL, July 2003. 

4. B. Bragg, C. Cerimele, R. Delventhal, J. Gamble, 

O. Hill, R. Kincade, D. Lee, W. Long, R. 

McHenry, G. McSwain, K. Nagy, M. Richardson, 

R. Ried, B. Roberts, C. Scott, D. Smith, “A Design 

Study for an Aeroassist Flight Experiment,” NASA 

Johnson Space Center, JSC-20593, June 1985. 

5. C. Cerimele, J. Gamble, “A Simplified Guidance 

Algorithm for Lifting Aeroassist Orbital Transfer 

Vehicles,” AIAA-85-0348, AIAA 23rd Aerospace 

Sciences Meeting, Reno, Nevada, January 1985. 

6. T. Snook, R. McHenry, “Monte Carlo Evaluation 

of Aerobraking Guidance Algorithms,” NASA 

Johnson Space Center, JSC-22432, January 1987. 

7. T. Ro, E. Queen, “Mars Aerocapture Terminal 

Point Guidance and Control,” AIAA-98-4571, 

AIAA Atmospheric Flight Mechanics Conference, 

Boston, MA, August 1998. 

8. L. Bryant, M. Tigges, D. Ives, “Analytic Drag 

Control for Precision Landing and Aerocapture,” 

AIAA-98-4572, AIAA Atmospheric Flight 

Mechanics Conference, Boston, MA, August 1998. 

9. J. Masciarelli, S. Rousseau, H. Fraysse, E. Perot, 

“An Analytic Aerocapture Guidance Algorithm for 

the Mars Sample Return Orbiter,” AIAA-2000-

4116, AIAA Atmospheric Flight Mechanics 

Conference, Denver, CO, August 2000. 

10. E. Perot, H. Fraysse, S. Rousseau, J.C. Berges, 

“Comparison of an Analytical Predictor Corrector 

and a Terminal Point Controller for the Mars 

Sample Return Aerocapture,” AAAF 14-67 26-29 

March 2001. 

11. S. Rousseau, E. Perot, C. Graves, J. Masciarelli, E. 

Queen, “Aerocapture Guidance Algorithm 

Comparison Campaign,” AIAA-2002-4822, 

AIAA/AAS Astrodynamics Specialist Conference, 

Monterey, CA, August 2002. 

12. C.G. Justus, A. Duvall, and D.L. Johnson, 

“Engineering-Level Model Atmospheres For Titan 

And Neptune,” AIAA-2003-4803, 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Huntsville, AL, July 2003. 

13. R. Haw, “Titan Approach Navigation for the Titan 

Aerocapture Orbiter,” AIAA-2003-4802, 

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference, Huntsville, AL, July 2003. 

14. E. Perot, S. Rousseau, “Importance of an On-Board 

Estimation of the Density Scale Height for Various 

Aerocapture Guidance Algorithms,” AIAA-2002-

4734, AIAA/AAS Astrodynamics Specialist 

Conference, Monterey, CA, August 2002. 
61

AEROCAPTURE SIMULATION AND PERFORMANCE FOR THE TITAN 

EXPLORER MISSION 

David W. Way 

Richard W. Powell 

Karl T. Edquist 

James P. Masciarelli 

Brett R. Starr 

ABSTRACT 

A systems study for a Titan aerocapture orbiter has been completed. The purpose of this study was to 

determine the feasibility and potential benefits of using aerocapture technologies for this destination. 

The Titan Explorer design reference mission is a follow-on to the Cassini/Huygens exploration of the 

Saturnian system that consists of both a lander and an orbiter. The orbiter uses aerocapture, a form of 

aeroassist, to replace an expensive orbit insertion maneuver with a single guided pass through the 

atmosphere. Key environmental assumptions addressed in this study include: the uncertainty in 

atmospheric density and high frequency atmospheric perturbations, approach navigation delivery 

errors, and vehicle aerodynamic uncertainty. The robustness of the system is evaluated through a 

Monte Carlo simulation. The Program to Optimize Simulated Trajectories is the basis for the 

simulation, though several Titan specific models were developed and implemented including: 

approach navigation, Titan atmosphere, hypersonic aeroshell aerodynamics, and aerocapture guidance. 

A navigation analysis identified the Saturn/Titan ephemeris error as major contributor to the delivery 

error. The Monte Carlo analysis verifies that a high-heritage, low L/D, aeroshell provides sufficient 

performance at a 6.5 km/s entry velocity using the Hybrid Predictor-corrector Aerocapture Scheme 

guidance. The current mission design demonstrates 3-sigma success without additional margin, 

assuming current ephemeris errors, and is therefore not dependent on the success of the 

Cassini/Huygens mission. However, additional margin above 3-sigma is expected along with the 

reduced ephemeris errors in the event of a successful Cassini mission. 

NOMENCLATURE 

BOC Beginning of Cassini 

c.g. Center of Gravity 

CAD Computer Aided Design 

CFD Computational Fluid Dynamics 

DoF Degree of Freedom 

EDL Entry, Descent, and Landing 

EOC End of Cassini 

GRAM Global Reference Atmospheric 

Model 

HYPAS Hybrid Predictor-corrector 

Aerocapture Scheme 

IRIS Infrared Interferometer Spectrometer 

L/D Lift to Drag Ratio 

LAURA Langley Aerodynamic Upwind 

Relaxation Algorithm 

MER Mars Exploration Rover 

MGS Mars Global Surveyor 

MPF Mars Pathfinder 

MRO Mars Reconnaissance Orbiter 

POST Program to Optimize Simulated 

Trajectories 

SEP Solar Electric Propulsion 

UVS Ultraviolet Spectrometer 

∆DOR Delta Differential One-way Ranging 

∆V Velocity Addition 

σ Standard Deviation 
62

BACKGROUND 

AEROCAPTURE DESCRIPTION 

Aerocapture, a form of aeroassist, is a propellantless alternative to the currently requisite allpropulsive planetary capture. Using drag to 

decelerate the vehicle, aerocapture replaces the 

expensive orbit insertion maneuver with a single 

guided pass through the atmosphere. To date, 

aerocapture has not been demonstrated in flight. 

In contrast, aerobraking uses many passes 

through the atmosphere to reduce the period of 

an elliptical orbit. This reduces, but does not 

eliminate, the propulsive capture requirement. 

Aerobraking has been used successfully in the 

Martian atmosphere by Mars Global Surveyor 

(MGS), and Mars Odyssey, and is planned for 

Mars Reconnaissance Orbiter (MRO). 

A nominal drag profile associated with the 

aerocapture pass is designed to remove all of the 

hyperbolic excess velocity and enough additional 

orbital energy to place the spacecraft in an 

elliptical orbit with the desired apoapsis. 

Because of the larger energy requirements, 

aerocapture occurs at altitudes much lower than 

aerobraking. A guidance system is used to target 

the desired exit conditions by reacting to changes 

in the atmosphere. Bank angle modulation is 

used to control the rate of ascent/descent, which 

indirectly affects the drag. The flight path angles 

required to fly full lift-up and full lift-down form 

a theoretical entry corridor. 

Figure 1 diagrams the sequence of aerocapture 

events. At the first apoapsis after the aerocapture 

pass, a small propulsive maneuver must be 

completed to raise the periapsis to the desired 

altitude. The periapsis must be raised during the 

first orbit in order to prevent the vehicle from reentering the atmosphere a second time. Another 

small propulsive burn is typically performed at 

periapsis to clean-up any residuals in the desired 

science orbit apoapsis. 

TITAN EXPLORER MISSION 

The Titan Explorer design reference mission is a 

follow-on to the Cassini/Huygens exploration of 

the Saturnian system that consists of both a 

lander and an orbiter.1

 Both spacecraft are 

launched together on a single Delta IV-class 

launch vehicle in 2010. Figure 2 shows a 

Computer Aided Design (CAD) model of the 

stack packaged in a 4 m launch fairing. A Solar 

Electric Propulsion (SEP) module and a single 

Venus gravity assist provide a 6.25 year 

interplanetary cruise to the Saturn system.2

 

Both the orbiter and the lander are initially 

targeted for a direct entry to Titan. Thirty days 

prior to arrival, the orbiter releases the lander and 

executes a divert maneuver to the desired 

aerocapture approach trajectory. 

The orbiter provides a telecom link for the lander 

during Entry, Descent, and Landing (EDL) then 

completes an aerocapture to the desired science 

orbit (a near-polar 1700 km circular orbit). 

Following aerocapture, the heatshield and 

backshell are jettisoned, and the orbiter begins a 

three-year science mission. 

STUDY GOALS 

A systems study for a Titan aerocapture orbiter 

has been completed as part of the NASA Inspace Propulsion Program.3

 The purpose of this 

study was to determine the feasibility and 

potential benefits of using aerocapture 

technologies for this destination.4,5 The products 

of this study are a reference mission, baseline 

systems definition, and technology requirements 

that may be used by scientists, systems 

engineers, technology developers, and mission 

managers in planning future missions. This 

study provides additional value over previous 

systems studies because of the higher fidelity of 

the analyses and environmental models that were 

employed. 

Key environmental assumptions, central to 

successful aerocapture, are addressed in this 

study. These assumptions include the 

uncertainty in atmospheric density, high 

Figure 1: Aerocapture into Circular Orbit 
63

frequency atmospheric variability, approach 

navigation delivery errors, and vehicle 

aerodynamic uncertainty. Aerocapture risk is 

mitigated by quantifying the atmospheric 

uncertainty based on all available measurements, 

designing the vehicle to provide adequate 

aerodynamic control authority, developing a 

robust guidance system, and incorporating 

sufficient margins. The robustness of the system 

is evaluated through Monte Carlo simulation. 

SIMULATION HERITAGE 

The heritage of the Monte Carlo simulation used 

in this study is based upon previous Langley 

Research Center work on many diverse planetary 

missions that involve phases of atmospheric 

flight. These missions include aerobraking 

orbiters: MGS, Mars Odyssey, and MRO 

(scheduled for launch in 2005); direct lander 

entries: Mars Pathfinder (MPF), Genesis, 

Stardust, Mars 2001 Surveyor Lander 

(cancelled), Mars Exploration Rovers (MER), 

and Mars Science Laboratory (MSL) (planned 

for launch in 2009); and aerocapture proposals: 

Mars Surveyor 2001 Orbiter (cancelled) and 

Mars Premier Orbiter (aerocapture option not 

adopted). The current simulation leverages this 

experience in atmospheric flight and applies it to 

a new destination, Saturn’s largest moon, Titan. 

SIMULATION DEVELOPMENT 

TRAJECTORY SIMULATION 

To aid in the systems study activity, a high 

fidelity Monte Carlo trajectory simulation has 

been developed to simulate flight through the 

Titan atmosphere during aerocapture. This 

simulation provides data and statistics used to 

quantify mission success probabilities, evaluate 

candidate guidance algorithms, and provide the 

technical feedback required for mission and 

aeroshell design (aerodynamic loads, maximum 

heat rate, integrated heat loads, orbit 

circularization fuel, etc.). 

The Program to Optimize Simulated Trajectories 

(POST) is the basis for this simulation.6

 

However, several Titan specific models were 

developed and implemented to support this 

work. These models include: approach 

navigation, Titan atmosphere, hypersonic 

aeroshell aerodynamics, and aerocapture 

guidance. These models are discussed in more 

detail. 

ATMOSPHERE 

An engineering-level atmosphere model, denoted 

Titan-GRAM, was developed for this study.7

 

Titan-GRAM is similar to and based upon the 

Mars Global Reference Atmosphere Model, 

Mars-GRAM, which has been used for in the 

design and operations support for many Mars 

exploration projects. 

Titan-GRAM atmospheric density predictions 

are based on minimum, nominal, and maximum 

density vs. altitude profiles predicted by Yelle et 

al.8

 The Yelle models are based on observations 

from Voyager 1 radio science, Infrared 

Interferometer-Spectrometer (IRIS), and 

Ultraviolet Spectrometer (UVS). Shown in 

Figure 3, the Yelle density profiles include 

density variation due to latitude, season, and 

diurnal effects as well as measurement 

uncertainty. 

SEP Prop

Module

Solar

Arrays

Orbiter

Lander

Figure 2: Launch Configuration 
64

Within Titan-GRAM, an atmospheric density 

control parameter, fminmax, is used to linearly 

interpolate between the Yelle profiles. An fminmax

of 1.0 corresponds to the maximum expected 

density for a given altitude, while fminmax of -1.0 

corresponds to the minimum expected density. 

A sinusoidal variation of fminmax with latitude was 

implemented to simulate latitudinal density 

gradients during an aerocapture pass. 

Within the trajectory simulation, fminmax is varied 

as a function of latitude to capture the expected 

latitudinal gradients. A perturbation model, 

based on gravity wave theory, is also included 

for use in the Monte Carlo analysis with a 

maximum perturbation (1σ) of 10% the mean 

density. Figure 4 shows a sample of perturbed 

density profiles generated by Titan-GRAM. 

AERODYNAMICS 

An aerodynamic model for the reference 

spacecraft has been developed using high-fidelity 

computations. The reference spacecraft has a 70 

deg sphere-cone heatshield, similar to the Viking 

Mars Lander entry vehicle, and a bi-conic 

backshell. The configuration is shown in Figure 

5. 

Constant normal and axial force aerodynamic 

coefficients are used for the aerocapture pass 

simulation and are based on Langley 

Aerodynamic Upwind Relaxation Algorithm 

(LAURA) Computational Fluid Dynamics 

(CFD) results in the hypersonic regime. 

LAURA solves the viscous fluid dynamic 

equations on a structured grid with built-in 

adaptation.9

 Thermal and chemical nonequilibrium models are used to calculate the 

high-temperature flowfield behind the bow 

shock. 

The high heritage, L/D = 0.25, aeroshell 

configuration provides 3.5 degrees of theoretical 

corridor width at a 6.5 km/sec entry velocity. A 

higher entry velocity of 10 km/s results in a 4.7 

degree theoretical corridor. 

Figure 3: Yelle et al. Titan Atmospheric Density Profiles 
65

GUIDANCE 

A Hybrid Predictor-corrector Aerocapture 

Scheme (HYPAS) aerocapture guidance 

algorithm was developed and included in the 

simulation.10 The HYPAS algorithm uses an 

analytic method, based on deceleration due to 

drag and altitude rate error feedback, to predict 

exit conditions and then adjust the bank angle 

command in order to achieve a target apoapsis 

altitude and orbit inclination at atmosphere exit. 

The HYPAS guidance consists of two phases: 

the “capture phase”, in which the guidance 

establishes pseudo-equilibrium glide conditions; 

and an “exit phase”, in which exit conditions are 

predicted, assuming a constant altitude rate, and 

the lift vector is adjusted to null the error 

between predicted and target apoapsis. Figure 6 

shows the guidance phases during an aerocapture 

pass. 

Bank reversals maintain inclination error within 

desired limits. All reference values are 

computed and updated during flight. The 

HYPAS algorithm was adapted for use at Titan, 

and two sets of guidance initialization 

parameters were developed: one for the 6.5 km/s 

entry, and one for the 10.0 km/s entry. Monte 

Carlo trajectory simulations were run with this 

guidance to determine overall aerocapture 

performance. 

A pseudo bank controller was developed to 

mimic the dynamics of a flight control system in 

a Three Degree-of-Freedom (3-DoF) simulation. 

These effects include a control system time lag 

and a finite system response, limited by a 

maximum angular acceleration and a maximum 

angular velocity. The bank angle controller 

analytically calculates the time required, and 

resulting angular travel necessary, to complete 

the maneuver to the commanded attitude. It has 

been found that including this type of controller 

in a 3-DoF simulation provides a good 

approximation to Six Degree-of-Freedom (6-

DoF) dynamics. 

Because the aerocapture spacecraft performs 

bank reversals to maintain inclination accuracy, 

and because these reversals could take as much 

as15 seconds to complete, the trajectory 

simulation must model the effects of an attitude 

controller. These bank reversals force the 

spacecraft off of the optimum flight profile that 

Figure 4: Sample Titan-GRAM Density Perturbations 

Access Opening

.762 m x .864 m

.635 m

.721 m

.740 m

3.75 m

Backshell

Heatshield

Figure 5: Aeroshell Configuration 
66

the guidance is trying to follow. Not including 

the error produced during this maneuver would 

result in overly optimistic conclusions regarding 

the vehicle’s targeting ability and the required 

circularization ∆V. 

NAVIGATION 

Initial states were provided by a JPL navigation 

assessment that assumed post-Cassini ephemeris 

knowledge and the following data sources: twoway Doppler and ranging, ∆DOR, and optical 

navigation.11 These assumptions resulted in a 3σ

delivery flight path angle dispersion of ±0.93 

deg. Figure 7 shows the delivery footprint in the 

B-plane. The dashed line in this figure is a 

radius vector to the nominal aim-point. 

The three dominant contributors to this delivery 

error were Saturn and Titan ephemeredes, 

maneuver execution error, and optical data 

measurement error. The current Cassini mission 

is expected to improve the ephemeris errors by a 

factor of six. However, this improved navigation 

is not guaranteed, but rather contingent upon the 

successful completion of the Cassini mission. 

Therefore, both Beginning of Cassini (BOC) and 

End of Cassini (EOC) states were evaluated. For 

the purposes of this study, it was assumed that 

the flight path angle dispersions would degrade 

by approximately 52% with the use of BOC 

states. 

Since Titan is a moon of Saturn with an orbital 

period of approximately 16 days, the mission 

designer has a wide choice in approach velocities 

for any mission opportunity (Titan’s velocity 

could either add or subtract from the nominal 

Saturn approach velocity). Intercepting Titan at 

different true anomalies easily tailors the entry 

velocity, with only small changes in the 

incoming hyperbolic approach trajectory. Entry 

velocities of 6.5 km/s and 10.0 km/s were 

considered. A 15% increase in flight path angle 

dispersions was levied on the higher entry 

velocity. 

Figure 7: Delivery Error in the B-plane 

1

2

3

4

5

6

Initial state

(Hyperbolic App)

Atmospheric entry interface

Periapsis

Atmospheric exit

Exit phase

(const. bank)

Capture phase

(const. bank)

Exit phase

(variable bank)

Capture phase

(variable bank)

Figure 6: Phases of HYPAS Aerocapture Guidance 
67

The Table 1 summarizes the entry flight path 

angle dispersions assumed in this study along 

with the theoretical corridor for an L/D = 0.25. 

Further navigational assessment is required to 

validate these assumptions. 

Table 1: Entry Flight Path Angle Uncertainties 

Entry 

Velocity 

EOC 

Ephemeris 

BOC 

Ephemeris 

Theoretical 

Corridor 

6.5 km/s ± 0.93 deg ± 1.42 deg 3.5 deg 

10 km/s ± 1.07 deg ± 1.63 deg 4.7 deg 

RESULTS 

MONTE CARLO ANALYSIS 

System performance, risk, and robustness are 

measured by generating statistics from Monte 

Carlo simulations of the Titan aerocapture. 

Many (generally 2000) individual aerocapture 

trajectories are simulated with random 

perturbations applied to initial entry conditions, 

vehicle aerodynamics, vehicle mass properties, 

and Titan atmospheric conditions. This flight 

simulation is composed of three main parts: a 

POST2 trajectory simulation, which integrates 

all of the models discussed above; an executive 

Monte Carlo script, which coordinates the 

generation and execution of 16 parallel 

simulations; and various supporting scripts for 

sampling random distributions, compiling and 

Table 2: Monte Carlo Uncertainties 

Category Variable Nominal ± 3σ or 

min/max 

Distribution 

Initial 

Conditions 

 

 x- position 603.3 km From 

covariance 

Correlated 

 y- position -390.8 km From 

covariance 

Correlated 

 z- position 3502 km From 

covariance 

Correlated 

 x- velocity -3.363 

km/s 

From 

covariance 

Correlated 

 y- velocity -4.123 

km/s 

From 

covariance 

Correlated 

 z- velocity -3.734 

km/s 

From 

covariance 

Correlated 

Atmosphere 

 Perturbation seed 1 1/9999 Uniform 

(integer) 

 Fminmax bias 0 -0.53/+0.53 Uniform 

Aerodynamics 

 Trim angle-ofattack -16 deg ± 2.0 deg Normal 

 CA (axial force) 1.48 ± 3% Normal 

 CN (normal force) -0.05 ± 5% Normal 

Mass Properties 

 Axial c.g.(Zcg/D) 0.1979 ± 0.00848 Normal 

 Radial c.g.(Xcg/D) 0.0231714 ± 0.00184 Normal 
68

formatting output data, evaluating metrics and 

statistics, and producing plots and figures. Table 

2 lists the uncertainties used in this study. 

BEGINNING OF CASSINI 

The first scenario examined is for a navigation 

ephemeris uncertainty consistent with knowledge 

prior to the Cassini mission. This combination 

of large navigation uncertainty and low entry 

velocity, 6.5 km/s, results in the most 

challenging conditions. Statistics for apoapsis 

altitude and circularization ∆V are presented in 

Table 3. 

Figure 8 shows the aerocapture corridor (flight 

path angle) as a function of fminmax (density). The 

theoretical corridor is bounded by the full liftdown and full lift-up cases. The plus (+) 

symbols show the range of fminmax during the 

active guidance portion of the aerocapture pass, 

due to latitudinal variation of fminmax. The circles 

indicate the fminmax at periapsis. The bias in the 

data towards the higher values of fminmax is again 

due to the latitudinal variation of fminmax, since 

the aerocapture pass occurs over northern 

latitudes. The target flight path angle was 

chosen to capture as many of the cases as 

possible into the theoretical corridor. The large 

spread in entry flight path angle, compared to the 

theoretical corridor, suggests small margins. 

Table 3 Performance Statistics, BOC 6.5 km/s 

 Apoapsis 

altitude 

(km) 

Circularization 

∆V (m/s) 

Minimum 1240.5 157.2 

Maximum 2166.3 293.3 

Mean 1691.3 179.6 

1σ ± 63.9 ± 11.1 

3σ ± 191.8 ± 33.2 

Figure 9 shows a histogram of the final apoapsis 

altitude. The mean apoapsis of 1691.3 km, with 

a standard deviation of +/- 63.9 km, compares 

well with the 1700 km target altitude. 

Additionally, the 3σ range of +/- 191.8 km is 

within the target range of < 200 km, which 

indicates that the guidance can be tuned to 

capture a large percentage of the theoretical 

corridor. Only one case (#505) failed to capture. 

However, this case represents a 4.5-sigma case 

for entry flight path angle, which has a 

probability of occurrence of only 1 in nearly 

15,000. 

Figure 10 shows the histogram for the required 

∆V. This ∆V includes both the periapsis raise 

maneuver and the final circularization burn. The 

3-sigma (99.86 %-tile) value is 212.9 m/s. 

END OF CASSINI 

These Monte Carlo results are representative of 

navigation ephemeris uncertainty post-Cassini 

for a 6.5 km entry velocity. Statistics for 

Figure 8: Entry Corridor, BOC 6.5 km/s 

Figure 9: Apoapsis Altitude, BOC 6.5 km/s 
69

apoapsis altitude and circularization ∆V are 

presented in Table 4. 

Figure 11 shows the aerocapture corridor (flight 

path angle) as a function of fminmax (density). 

The effect of the improved (reduced) uncertainty 

expected from the Cassini mission is evident by 

the tighter grouping in flight path angle as 

compared to Figure 8. The size of the flight path 

angle dispersions, compared to the theoretic 

corridor, suggests increased margins for this 

scenario. Because of the arrival geometry, the 

aerocapture pass occurs over northern latitudes – 

entering over the northern pole and exiting near 

the equator. Therefore, the mean fminmax (~0.3) is 

positive. The target flight path angle was chosen 

to bring this mean to the center of the theoretical 

corridor. 

Table 4 Performance Statistics, EOC 6.5 km/s 

 Apoapsis 

altitude 

(km) 

Circularization 

∆V (m/s) 

Minimum 1327.8 156.0 

Maximum 2196.6 252.8 

Mean 1697.7 177.7 

1σ ± 63.5 ± 9.5 

3σ ± 190.4 ± 28.6 

Figure 12 shows a histogram of the final 

apoapsis altitude. The mean apoapsis is 

1697.7 km, with a standard deviation of 63.5 

km. The 3-sigma range of +/- 190.4 km is 

within the target range of < 200 km. All 

cases captured within 500 km of the target 

apoapsis. 

Figure 13 shows the histogram and statistics 

for the required ∆V. The 3-sigma (99.86 %-

tile) value is 206.3 m/s. The maximum ∆V 

was 252.8 m/s. 

Figure 10: Circularization ∆V, BOC 6.5 km/s Figure 11: Entry Corridor, EOC 6.5 km/s 

Figure 12: Apoapsis Altitude, EOC 6.5 km/s 
70

CONCLUSIONS 

1. The JPL navigation analysis identified the 

Saturn/Titan ephemeris error as major 

contributor the delivery error. The current 

mission design demonstrates 3-sigma 

success, without additional margin, 

assuming BOC ephemeris errors, and is 

therefore not dependent on the success of 

the Cassini/Huygens mission. However, 

additional margin above 3-sigma is expected 

along with the reduced EOC ephemeris 

errors in the event of a successful Cassini 

mission. 

2. Uncertainty in the Titan atmospheric 

density, including high frequency 

perturbations, is the single largest unknown. 

To mitigate this risk, sufficient margin and 

conservatism are carried in the design of the 

entry conditions, aeroshell, and guidance 

system. While arrival during a particular 

season would reduce the expected density 

range, the full density range was used in the 

Monte Carlo analysis. 

3. The Monte Carlo analysis verifies that a 

high-heritage, low L/D, aeroshell provides 

sufficient performance at a 6.5 km/s entry 

velocity. A mid L/D aeroshell technology 

development is not required. Additional 

aerocapture performance is also available at 

higher entry velocities, 10 km/s, but must be 

traded against increased Thermal Protection 

System requirements. 

4. The Monte Carlo analysis demonstrates that 

the HYPAS guidance is robust and provides 

acceptable performance. Approximately 

92% of the theoretical corridor is captured 

using this algorithm while requiring only 

slightly more than 200 m/s of on-orbit ∆V to 

achieve the target science orbit. 

REFERENCES 

1. Bailey, R., Hall, J., and Spilker, T., “Titan 

Aerocapture Mission and Spacecraft Design 

Overview,” AIAA-2003-4800, Conference 

Proceedings of the 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Huntsville, AL, 

July 2003. 

2. Noca, M., Bailey, R., and Dyke, R., “Titan 

Explorer Mission Trades from the 

Perspective of Aerocapture,” AIAA-2003-

4801, Conference Proceedings of the 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Huntsville, AL, 

July 2003. 

3. James, B. and Munk, M., “Aerocapture 

Technology Development Within the NASA 

In-Space Propulsion Program,” AIAA-2003-

4654, Conference Proceedings of the 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Huntsville, AL, 

July 2003. 

4. Hall, J., Noca, M., and Bailey, R., “Cost -

Benefit Analysis of the Aerocapture Mission 

Set,” AIAA-2003-4658, Conference 

Proceedings of the 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Huntsville, AL, 

July 2003. 

5. Lockwood, M. K., “Titan Aerocapture 

Systems Analysis,” AIAA-2003-4799, 

Conference Proceedings of the 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Huntsville, AL, 

July 2003. 

Figure 13: Circularization ∆V, EOC 6.5 km/s 
71

6. Bauer, G. L., Cornick, D. E., and Stevenson, 

R., “Capabilities and Applications of the 

Program to Optimize Simulated Trajectories 

(POST),” NASA CR-2770, February 1977. 

7. Justus, C. and Duvall, A., “EngineeringLevel Model Atmospheres for Titan and 

Neptune,” AIAA-2003-4803, Conference 

Proceedings of the 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Huntsville, AL, 

July 2003. 

8. Yelle, R. V., Strobel, D. F., Lellouch, E., 

and Gautier, D., “Engineering Models for 

Titan's Atmosphere,” in ESA report SP1177, “Huygens Science, Payload and 

Mission,” August 1997. 

9. Cheatwood, F. M. and Gnoffo, P. A., 

“User’s Manual for the Langley 

Aerodynamic Upwind Relaxation Algorithm 

(LAURA),” NASA TM-4674, April 1996. 

10. Masciarelli, J. and Queen, E., “Guidance 

Algorithms for Aerocapture at Titan,” 

AIAA-2003-4804, Conference Proceedings 

of the 39th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference and Exhibit, 

Huntsville, AL, July 2003. 

11. Haw, R., “Approach Navigation for a Titan 

Aerocapture Orbiter,” AIAA-2003-4802, 

Conference Proceedings of the 39th

AIAA/ASME/SAE/ASEE Joint Propulsion 

Conference and Exhibit, Huntsville, AL, 

July 2003. 
72 

PRELIMINARY AEROTHERMODYNAMICS OF TITAN AEROCAPTURE 

AEROSHELL 

Naruhisa Takashima

AMA Inc. 

Hampton, VA 23666 

Brian R. Hollis and E. Vincent Zoby 

NASA Langley Research Center 

Hampton, VA 23681 

Kenneth Sutton 

National Institute of Aerospace 

Hampton, VA 23666 

Joe Olejniczak 

NASA Ames Research Center 

Moffett Field, CA 94035 

Michael J. Wright and Dinesh Prabhu 

ELORET Corporation 

Sunnyvale, CA 94087 

ABSTRACT 

Aeroheating environments for a Titan aerocapture mission were computed for the Titan Aerocapture Systems 

Analysis study funded through the In-Space Propulsion Program Office at NASA Marshall Space Flight Center. In 

this work, the convective heating environments for a candidate Titan probe (70 deg. sphere-cone geometry) are 

presented. The sensitivities of these environments to the computational grids, mass diffusion models, and reaction 

rates are examined in the context of axisymmetric flow using two different flow solvers. The lessons learned are 

applied to the forebody of the probe at an angle of attack of 16 deg. Results of computations from two different flow 

solvers are compared to reduce uncertainties in the predicted aerothermal environments. At the peak convective 

heating condition, the convective heating rate at the stagnation point is 28 W/cm2

 and the maximum heating rate, 

which occurs at the nose of the axisymmetric aeroshell is 46 W/cm2

. A smooth-wall transition criterion of Reθ of 

200 is used in determining the likelihood of flow transition on the forebody and shown to occur on the leeside. 

Turbulent flow calculations show that the level of heating on the leeside increases by 250%, and the location of the 

maximum heating point moves from the nose to the leeside shoulder. Finally, aerodynamic force coefficients 

extracted from the computed solutions are shown to be in good agreement with those assumed in the generation of 

flight trajectories. The three-dimensional aerothermal environments, along with the predicted radiative heating 

environments, are used in the selection and sizing of the Thermal Protection System for the forebody of the probe. 

NOMENCLATURE

Alt Altitude (km) 

AoA Angle of attack (deg) 

CL Lift coefficient 

CD Drag coefficient 

D Reference diameter (3.75 m) 

L/D Lift-to-Drag Ratio 

Mach Mach number 

P Pressure (Pa) 

Reθ Momentum thickness Reynolds 

number. 

ρ and Rho Density (kg/m3

) 

T Translational temperature (K) 

Tv Vibrational temperature (K) 

TPS Thermal Protection System 

V Velocity (m/s) 
73 

INTRODUCTION

Titan, the largest moon of Saturn, is the only 

known moon in the solar system that has a fully 

developed atmosphere. The density of the atmosphere 

of Titan is actually greater than that of the Earth with 

a surface pressure 50% greater than that of Earth. The 

atmosphere is composed mainly of nitrogen, which 

accounts for approximately 94% by volume, with 

balance of argon and significant traces of 

hydrocarbon elements. Hydrocarbon elements are the 

building blocks for amino acids, a necessary 

ingredient for formation of life. For this reason, Titan 

is of significant interest to the scientific community 

for the understanding of early formation of life here 

on Earth. Voyager 1 was the first mission to Titan 

taking pictures of its dense atmosphere. Currently, 

the Cassini spacecraft with Huygens probe on board 

is en route to Titan. The joint NASA/ESA mission 

expects the spacecraft to arrive at Saturn on July 1st

, 

2004 and release its probe on December 25th, 2004. 

As a potential follow-on mission to the CassiniHuygens mission, a systems analysis study of an 

aerocapture mission to Titan was initiated by the 

NASA In-Space Propulsion Program Office at the 

NASA Marshall Space Flight Center (MSFC). Past 

studies have shown a vehicle mass saving of up to 

66% can be realized using aerocapture technology 

compared to an all-propulsive mission to Titan. 

However, by design, these past analyses were of low 

fidelity and many assumptions were made to reach 

the conclusions. Hence, the goals for the present 

study include providing higher fidelity systems 

definition and sensitivities for Titan aerocapture 

mission that can be used by scientists, mission 

planners, technology planners, technologists and 

future mission managers; and to perform higher 

fidelity analyses for validating and updating previous 

assumptions. The system analysis study was 

performed by a multi-center NASA team with 

participation from NASA Ames Research Center 

(ARC), Jet Propulsion Laboratory, Johnson Space 

Center, Marshall Space Flight Center, and led by 

Langley Research Center (LaRC).1

 

This paper will present the results of the 

computations performed by LaRC and ARC to 

determine the aeroheating environments and will 

focus on the convective heating portion of these 

environments. A companion paper will present the 

results of the radiative heating analyses.2

 Prior to the 

work, extensive analyses, both numerical and 

experimental, were performed to determine the 

aeroheating environment of the Huygens probe.3,4,5,6

Analyses have shown that significant radiative 

emission of mainly CN violet system can occur 

within the shock layer due to non-equilibrium 

condition, and the amount of the CN radiation is 

sensitive to the freestream gas composition. For the 

worst case condition, the level of radiative heat flux 

was determined to be twice that of the convective 

heating flux.4 

TITAN AEROCAPTURE REFERENCE 

CONCEPT

The forebody design of the aeroshell used by the 

study is shown in Fig. 1. The design is based on the 

Mars Smart Lander (MSL) aeroshell design,7

 which 

has a 70 deg. sphere cone forebody and a bi-conic 

aftbody. For the Titan aeroshell, the MSL design was 

geometrically scaled to a diameter of 3.75 m, and the 

aftbody shape was altered to meet the packaging 

requirements. The overall length of the vehicle is 

2.096 m, and the center of gravity is located such that 

an L/D of 0.25 is achieved at a trim angle of attack of 

16 deg. The study defined six reference trajectories 

for an entry speed of 6.5 km/s with a ballistic 

coefficient of 90 kg/m2

. Trajectories were created for 

combinations of three atmospheric density profiles: 

minimum, nominal and maximum, with lift vector 

directions of up (undershoot) and down (overshoot). 

The three density profiles are shown in Fig. 2, which 

were obtained with Titan atmosphere model, 

TitanGRAM, developed by Justus and Duvall.8

 The 

reference trajectories were computed using the 

Program for Optimization of Simulated Trajectories 

(POST).9

 

For the present study, aerothermodynamic 

analyses were performed primarily for conditions 

along the minimum atmosphere lift-up trajectory. 

Engineering analyses showed that the maximum 

stagnation point convective heating rate on a sphere 

occurs along the minimum atmosphere lift-up 

trajectory. Fig. 3 shows the points along the reference 

trajectory that the computations were performed and 

Table 1 lists the freestream at each trajectory point. 

Based on equilibrium engineering analysis, cases 3 

(t=253 s), 4 (t=281 s) and 5 (t=308 s) were predicted 

to be the peak convective heating, the peak dynamic 

pressure and the peak radiative heating, respectively. 
74 

There are many uncertainties in the composition 

of the Titan atmosphere.8

 The relative amounts of 

nitrogen, argon and methane will impact the overall 

heating rate on the vehicle. For radiative heating 

conservatism, the atmospheric composition was 

chosen as 95% N2 and 5% CH4 by volume along the 

entire reference trajectory. 

-2

-1.5

-1

-0.5

0

0.5

1

1.5

2

X

0 -0.5 -1 -1.5 -2

Z

01

Y

70 deg

.231 m

.445 m

65 deg

.231 m + 1.413 m = 1.644 m

Veh Diameter = 3.288 m

14% decrease from 3.75 m

HGA Diam = Diam at Break

.231 m + 2.4/2 m = 1.43 m

Veh Diameter = 2.862 m

24% decrease from 3.75 m

.296 m + 1.413 m = 1.709 m

Veh Diameter = 3.418 m

9% decrease from 3.75 m

HGA Diam = Diam at Break

.296 m + 2.4/2 m = 1.496 m

Veh Diameter = 2.992 m

20% decrease from 3.75 m

70 deg

.231 m

.445 m

65 deg

.231 m + 1.413 m = 1.644 m

Veh Diameter = 3.288 m

14% decrease from 3.75 m

HGA Diam = Diam at Break

.231 m + 2.4/2 m = 1.43 m

Veh Diameter = 2.862 m

24% decrease from 3.75 m

.296 m + 1.413 m = 1.709 m

Veh Diameter = 3.418 m

9% decrease from 3.75 m

HGA Diam = Diam at Break

.296 m + 2.4/2 m = 1.496 m

Veh Diameter = 2.992 m

20% decrease from 3.75 m

Fig. 1. Titan aeroshell design. 

 

Fig. 2. Titan atmosphere density profiles. 

0

100

200

300

400

500

600

700

800

0

1,000

2,000

3,000

4,000

5,000

6,000

7,000

8,000

0 100 200 300 400 500 600

Alt

(km)

V

(m/s)

Time (sec)

1

2

3

4

5

6

7

1

2

3

4 5 6

Fig. 3. Minimum atmosphere lift-up aeroheating 

reference trajectory. 

Table 1. Freestream conditions along minimum atmosphere lift-up reference trajectory. 

CASE Time Alt V AoA rho P T Mach

(sec) (km) (m/s) (deg) (kg/m^3) (Pa) (K)

1 174 420 6558 16 4.697E-06 0.188 131.9 27.8

2 222 316 6348 16 5.070E-05 2.293 149.9 25.3

3 253 269 5761 16 1.491E-04 6.910 152.7 22.7

4 281 245 4859 16 2.665E-04 12.350 152.6 19.2

5 308 237 3978 16 3.305E-04 15.250 152.4 15.7

6 339 238 3239 16 3.212E-04 14.840 152.4 12.8

7 397 255 2521 16 2.138E-04 9.893 152.7 9.9
75 

COMPUTATIONAL ANALYSES

As mentioned previously, the aeroheating 

environment computations were performed 

simultaneously during the course of the study by 

two teams from NASA Langley and Ames 

Research centers. Each team performed the 

calculations independently using different 

Computational Fluid Dynamics (CFD) programs. 

Once a set of calculations was made, the results 

were compared and analyzed by the two teams. 

If differences were observed, the causes were 

investigated and analyzed to the satisfaction of 

both teams. The purpose of having two totally 

independent parties calculating identical 

problems was to minimize the uncertainties in 

the results and gain greater confidence in the 

level of physical and numerical modeling in the 

codes. 

During the course of the study, computational 

analyses were performed in two stages. During the 

first stage, computations were performed on the 

forebody at a zero degree angle of attack, i.e., 

axisymmetric flow, for the peak convective heating 

condition (Case 3) to perform sensitivity analyses 

with respect to transport/kinetic models. After the 

completion of the first stage, the second stage 

involved computing three-dimensional flow on the 

forebody for all seven trajectory points. The results 

from the second stage were delivered to the TPS team 

for TPS sizing purpose. 

Sensitivity analyses were performed using two 

different CFD codes: Langley Aerothermodynamic 

Upwind Relaxation Algorithm (LAURA),10

developed at LaRC, and the Data Parallel Line 

Relaxation (DPLR) program,11 developed at ARC. 

Both CFD codes solve the three-dimensional NavierStokes equation using finite volume method with 

finite rate chemistry and thermal nonequilibrium. 

LAURA uses Roe’s flux difference splitting with 

Yee’s second-order symmetric total variation 

diminishing scheme to model the inviscid fluxes. 

Steady state solution is obtained using either point or 

line relaxation. The code supports MPI with multiblock structured grid to decrease computing time. 

DPLR uses third-order modified Steger-Warming 

flux vector splitting to calculate the inviscid fluxes 

and data-parallel line relaxation method to reduce the 

time to steady state convergence. Both codes use 

second order central differencing for the viscous 

fluxes and turbulent flow is modeled with the 

Baldwin-Lomax turbulence model. Both codes have 

been widely used in predicting aerothermal 

environments for several planetary probes.7,10,12 Each 

code was executed by its respective center during the 

course of the study. 

RESULTS

SENSITIVITY ANALYSES

To gain insight and reduce the uncertainties in 

the numerical results, sensitivity analysis on the 

diffusion model and chemical kinetic rates were 

made by performing calculations for the peak 

convective heating condition (Case 3) at zero degree 

of angle attack. For all the runs, the codes were 

executed with finite rate chemistry and thermal 

nonequilibrium using Park’s two-temperature 

model.13 For the nominal case, a total of 18 species 

was assumed to be present in the gas mixture: N, N+

, 

N2, N2

+

, C, C+, C2, CH, CH2, CH3, CH4, H, H+, H2, 

CN, CN+, NH and e-. “Super catalytic” surface 

boundary condition was chosen where the species 

fraction is fixed at freestream values of 95% N2 and 

5% CH4 at the wall. Forcing the gas mixture to 

recombine to the freestream composition causes the 

chemical energy to be released back into the flow, 

which results in conservative wall heating rates. The 

wall temperature was assumed to be at radiative 

equilibrium with surface emissitivity of 0.90, and the 

boundary layer was assumed to be laminar. 

Fig. 4. shows the grid and the temperature 

contour for the axisymmetric solution using LAURA. 

The grid for the solution was obtained using the builtin grid alignment capability of LAURA. The 

capability allows for grid clustering within the 

boundary layer and near the shock, as well as, 

adaptation of the outer boundary to the bow shock. 

The degree of adaptation is controlled through userdefined parameters. 

Proper cell spacing near the wall is essential for 

calculating accurate heating rates. The wall cell 

spacing in LAURA is controlled by specifying the 

wall cell Reynolds number: 

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ ∆ = μ

ρ a η Rew ( 1 ) 

where a is the local sound speed, ∆η is the cell height 

at the wall and μ is the local viscosity. Grid 

sensitivity analysis showed that, as seen in Fig. 5, 

Rew = 10 is sufficient to resolve the laminar boundary 

heating rate. Requirements for DPLR are similar. The 

LAURA solutions presented in this work was 

obtained using Rew = 1. 
76 

Fig. 6 shows comparison of convective heating 

rates along the surface using different diffusion 

model and chemical species. 12 species mixture is 

simply the 18 species mixture mentioned above 

without the ions and electron. The figure shows a 

variation of approximately 25% in stagnation point 

heating rate depending on the diffusion model and 

the chemical species. The two LAURA solutions 

with 18 species show that a higher stagnation heating 

rate of 50 W/cm2

 is predicted with the multicomponent diffusion model, compared to the 

stagnation heating rate of 42 W/cm2

 computed with 

constant Schmidt number. Note the good agreement 

in the heating rate distribution between the two codes 

when the same diffusion model of constant Schmidt 

number of 0.5 and the total number of species are 

used. 

Fig. 7 shows the variation of the translational 

temperature profile along the stagnation streamline 

calculated for the nominal case by DPLR and 

LAURA with three different chemical kinetic 

models. The forward reaction rate coefficients for 

each model are listed in Table 2. The backward 

reaction rates were calculated with equilibrium 

constants for both codes. Although all the solutions 

shown predicted the same stagnation point heating 

rate of approximately 50 W/cm2

, the plot shows that 

the shock stand off distance and the peak post shock 

temperature vary with different chemical kinetic 

models. The profile labeled LAURA (Nelson ‘91) 

was computed with LAURA, assuming heavy 

particles as the collision partner for the ionization 

reactions, as listed by Nelson.3

 During the study, it 

was confirmed through a private conversation with 

the author that the ionization reactions in the article 

were not appropriate, and that the collision partner 

should be electrons and not heavy particles. The 

profile labeled LAURA (Park ‘85) was computed 

with the same reaction rates as Nelson ‘91 but with 

the correct electron ionization reactions. Comparison 

of the two profiles shows that the heavy particle 

ionization reactions contribute to a decrease in both 

the post shock temperature and shock stand off 

distance, which is caused by the greater ionization of 

particles. Comparison of two profiles labeled 

LAURA (Park ’85) and LAURA (Park ’89), which 

are taken from rates published by Chul Park,13,14

shows the influence of nitrogen dissociation rates on 

the shock stand off and temperature. The faster 

nitrogen dissociation rates for Park ’89, as shown in 

Table 2, causes greater dissociation of nitrogen 

molecules. This decreases the energy within the 

shock layer that results in slightly lower post shock 

temperature and shock stand off distance. Lastly, as 

with the diffusion model comparison, when the same 

chemical kinetic model Park ’85 is used, the 

stagnation line temperature profiles calculated by the 

two codes are in relatively good agreement. 

 Table 3 lists the convective and radiative 

heating rates at the axisymmetric stagnation point 

calculated using the three chemical kinetic models. 

The radiative heating rates were calculated using 

LORAN15 with the flow field results from the 

LAURA calculations. The existence of Cyano 

radical, CN, contributes to the large values of 

radiative heating rates, which are approximately 

twice the amount of convective heating rates. The 

table shows that the convective heating rates are 

insensitive to the kinetic models, whereas, the 

radiative heating rates vary by approximately 225%. 

The results indicate that, as expected, the radiative 

heating rate is sensitive to the different post shock 

temperature and shock stand off distance computed 

by the different models. A difference in shock stand 

off distance of just 1 cm can result in a radiative 

heating rate difference of approximately 30%. 

Similar results were calculated using NEQAIR16 and 

DPLR. Details of the radiation calculations are 

presented in the companion paper.2 

X

Z

-2 -1 0 1 2

-1.5

-1

-0.5

0

0.5

1

9545.48

9177.13

8808.78

8440.44

8072.09

7703.74

7335.4

6967.05

6598.7

6230.36

5862.01

5493.67

5125.32

4756.97

4388.63

4020.28

3651.93

3283.59

2915.24

2546.9

2178.55

1810.2

1441.86

1073.51

705.164

336.817

T (K)

Fig. 4. Temperature contour at peak heating 

condition. 
77 

Table 2. Forward reaction rates. 

Reaction Forward reaction rate coefficient kf [cm3

/mole-s] 

Dissociation Reactions Ta= (T Tv)

0.5

Nelson ’91, Park ’89,Park ‘85

 C2 + M → C + C + M 9.68 × 1022 Ta

-2.0 exp( -71,000 /Ta ) 

 CH + M → C + H + M 1.13 × 1019 Ta

-1.0 exp( -40,193 /Ta ) 

 CN + M → C + N + M 1.00 × 1023 Ta

-2.0 exp( -90,000 /Ta ) 

 CH4 + M → CH3 + H + M 2.25 × 1027 Ta

-1.87 exp( -52,900 /Ta ) 

 CH3 + M → CH2 + H + M 2.25 × 1027 Ta

-1.87 exp( -54,470 /Ta ) 

 CH2 + M → CH+ H + M 2.25 × 1027 Ta

-1.87 exp( -50,590 /Ta ) 

 NH + M → N + H + M 1.13 × 1019 Ta

-1.0 exp( -41,820 /Ta ) 

 H2 + M → H + H + M 1.47 × 1019 Ta

-1.23 exp( -51,950 /Ta ) 

 

Nelson ’91, Park ‘85 

 N2 + M → N + N + M 3.70 × 1021 Ta

-1.6 exp( -113,200 /Ta ) 

 

Park ‘89 

 N2 + N2 → N + N + N2 7.0 × 1021 Ta

-1.6 exp( -113,200 /Ta ) 

 N2 + N → N + N + N 3.0 × 1022 Ta

-1.6 exp( -113,200 /Ta ) 

Exchange Reactions ( Ta = T ) 

Nelson ’91, Park ’89,Park ‘85 

 C + N2 → CN + N 1.11 × 1014 Ta

-0.11 exp( -23,000 /Ta ) 

 CN + C→ C2 + N 3.00 × 1014 Ta

0.0 exp( -18,120 /Ta ) 

 C2 + N2 → CN + CN 7.10 × 1013 Ta

0.0 exp( -5,330 /Ta ) 

 H + N2 → NH + N 2.20 × 1014 Ta

0.0 exp( -71,370 /Ta ) 

 H2 + C → CH + H 1.80 × 1014 Ta

0.0 exp( -11,490 /Ta ) 

 CN+ + N → CN + N+

 9.80 × 1012 Ta

0.0 exp( -40,700 /Ta ) 

 C+ N→ CN+

 + e- 1.00 × 1015 Ta

1.5 exp( -164,400 /Ta ) 

 C+ + N2 → N2

+

 + C 1.11 × 1014 Ta

-0.11 exp( -50,000 /Ta ) 

Associative Ionization Reactions ( Ta = T ) 

 N + N → N2

+

 + e- 1.79 × 1009 Ta

0.77 exp( -67,500 /Ta ) 

Ionization Reactions (Ta = Tv) 

Nelson ‘91 

 N + M → N+

 + e- + M 2.50 × 1034 Ta

-3.82 exp( -168,600 /Ta ) 

 C + M → C+

 + e- + M 3.90 × 1033 Ta

-3.78 exp( -130,000 /Ta ) 

 H + M → H+

 + e- + M 5.90 × 1037 Ta

-4.0 exp( -157,800 /Ta ) 

 Ar + M → Ar+

 + e- + M 2.50 × 1034 Ta

-3.82 exp( -181,700 /Ta ) 

Park ’85, Park ‘89 

 N + e- → N+

 + e- + e- 2.50 × 1034 Ta

-3.82 exp( -168,600 /Ta ) 

 C + e- → C+

 + e- + e- 3.90 × 1033 Ta

-3.78 exp( -130,000 /Ta ) 

 H + e- → H+

 + e- + e- 5.90 × 1037 Ta

-4.0 exp( -157,800 /Ta ) 

 Ar + e- → Ar+

 + e- + e- 2.50 × 1034 Ta

-3.82 exp( -181,700 /Ta ) 
78 

X

qw (W/cm2

)

0 0.5 1 1.5 0

10

20

30

40

50

60

Tw = Radiative Equilibrium (Rew = 10.0)

Tw = Radiative Equlibrium (Rew= 1.0)

Peak Heating Condition (Case 3)

V∞ = 5761.3 m/s , ρ∞= 0.1491E-4 kg/m3

, T∞= 152 K

Laminar BL, Super Catalytic

Fig. 5. Wall cell Reynolds number sensitivity. 

X (m)

qw (W/cm2

)

0 0.5 1 1.5 0

10

20

30

40

50

LAURA ( 12 species, Constant Schmidt Number Sc=0.5)

DPLR (12 species, Constant Schmidt Number Sc=0.5)

LAURA (18 species, Constant Schmidt Number Sc= 0.5)

LAURA (18 species, Multi Component Diffusion)

Case 3: Peak Convective Heating

V∞ = 5761.3 m/s , ρ∞= 0.1491E-4 kg/m3

, T∞= 152.7 K

α = 0 deg. (Axisymmetric)

Laminar BL, Tw= Radiative Equilibrium, Super Catalytic

Fig. 6. Variation of convective heating rate along the 

forebody surface. 

Z (m)

T (K)

-0.15 -0.1 -0.05 0 0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

12000

DPLR (Park '85)

LAURA (Park '85)

LAURA (Park '89)

LAURA (Nelson '91)

Fig. 7. Translational temperature profile along the 

stagnation streamline. 

Table 3. Axisymmetric stagnation point heating rates for peak heating condition. 

 Convective (W/cm2) Radiative (W/cm2) 

LAURA (Nelson ’91) 49.32 

Figure 1. 60.93 

LAURA (Park ’89) 51 107.81 

LAURA (Park ’85) 51.86 138.52 
79 

THREE –DIMENSIONAL RESULTS

Three-dimensional calculations were performed 

along the minimum atmosphere lift-up trajectory for 

the entry speed of 6.5 km/s using LAURA with 18 

chemical species. Multi-component diffusion model 

with Park ’89 chemical kinetic rates were used for the 

calculation. Fig. 8 shows the multi-block structured 

grid used for the LAURA calculations. 

Fig. 9 shows the shock structure in the pitch 

plane for the freestream conditions along the 6.5 km/s 

reference trajectory listed in Table 1. All 

computations were made for an angle of attack of 16 

deg., which is the design trim angle for the aeroshell. 

Fig. 10 shows the variation of convective heating rate 

along the surface in the pitch plane for all seven 

trajectory points, and Fig. 11 shows the variation of 

convective heating rate at the nose along the 

trajectory. The maximum heating rate along the 

trajectory is 46 W/cm2

, which transpire at the Case 3 

peak heating condition. For all cases, the maximum 

heating rate occurs not at the flow stagnation point, 

which is located approximately at X= -.98 m, but at 

the windward side shoulder and the nose/geometric 

stagnation point. Fig. 12 shows the variation of Reθ

along the forebody pitch plane for all seven cases. 

Based on a conservative engineering transition 

criterion of Reθ of 200, the flow will likely transition 

to turbulent on the leeward side of the forebody some 

time before t=253 s (Case 3) and remain turbulent 

until after t=339 s (Case 6). 

Fig. 13 shows variation of convective heating 

rate along the forebody pitch plane for fully laminar 

and turbulent flows for Cases 3 and 5. Turbulent flow 

was modeled using the algebraic turbulence model of 

Baldwin and Lomax.17 The figure shows that the 

turbulent heating rate is 2.5 times that of the laminar 

heating rate on the leeward side, and the location of 

maximum heating rate has shifted from the nose and 

the windward side shoulder to the leeward side 

shoulder. Although the engineering criterion predicts 

the flow will be laminar on the windward side, a 10% 

increase in heating level is observed due to turbulent 

flow for both cases. 

X

Y Z 70o Sphere Cone (Resized MSL)

Rnose= 0.9124 m, Rshoulder= 0.09144 m

D= 3.75 m

16 Computational Blocks:

Nose Region: 4 Blocks (10 x 10 x 64)

Cone Region: 6 Blocks (20 x 10 x 64)

Shoulder Region: 6 Blocks (28 x 10 x 64)

Rew= 1.0

Freestream Velocity

Symmetry Plane

Fig. 8. Computational grid for the forebody 

calculation

Fig. 9. Shock structure in the pitch plane. 
80 

Table 4 lists the lift and drag coefficient, as well 

as the lift-to-drag ratio of the forebody at each 

trajectory points. Case 1 data is excluded from the 

table based on the finding that condition is above the 

limits for continuum Navier-Stokes calculation. The 

table shows that the configuration meets the 

preliminary L/D design requirement of 0.25 along the 

trajectory. The CG location at the 16° trim angle of 

attack is at z/D = 0.3 and x/D = 0.0197. 

Fig. 10. Variation of convective heating rate on the 

forebody for the seven trajectory points. 

0

10

20

30

40

50

150 200 250 300 350 400

Convective heating at nose

q (W/cm2)

Time (s)

Fig. 11. Variation of convective heating rate along 

the reference trajectory.

Fig. 12. Variation of Reθ along the forebody for each 

trajectory point. 

Fig. 13. Turbulent and laminar convective heating 

rates for Case 3 and 5.

Table 4. Vehicle aerodynamics. 

 Case 1 Case 2 Case 3 Case 4 Case 5 Case 6 Case 7 

CL * 0.360 0.360 0.357 

Figure 2. 0.357

0.360 0.362 

CD * 1.446 1.437 1.421 1.420 1.427 1.429 

L/D * 0.249 0.251 0.252 0.251 0.252 0.254 
81 

SUMMARY AND CONCLUSIONS 

Aeroheating environments along a reference 

trajectory were calculated for the Titan Aerocapture 

System Analysis study. The environment was 

calculated using two different CFD codes by two 

teams from different NASA Research centers for the 

purpose of reducing uncertainties through direct 

comparison of results. Sensitivity analyses showed 

that two codes are in good agreement when identical 

transport/chemical kinetic models are used. 

Axisymmetric flow calculations showed that while 

insensitive to the chemical kinetic rates, convective 

heating rates are sensitive to the diffusion model and 

the choice of chemical species. In contrast, the 

radiative heating rates calculated using the CFD 

solutions are sensitive to the shock stand off distance 

and the post shock temperature profile, which are 

influenced by the chemical kinetic model as 

expected. Radiation calculations using the CFD 

results showed that the majority of the heating on the 

aeroshell will be due to radiative heating, which will 

drive not only the TPS sizing but also the material 

selection due to the nature of radiative heating. For 

this reason, careful reexaminations of the chemistry 

models are needed to reduce the uncertainties in the 

radiative heating. 

Three dimensional forebody calculations showed that 

the reference configuration meets the design L/D 

requirement of 0.25 along the reference trajectory. 

Momentum thickness Reynolds number along the 

forebody indicates that the flow is likely to transition 

to turbulent on the leeward side. Turbulent 

calculations show that the convective heating level on 

the leeside is increased by 250%, and the location of 

the maximum heating moves from the nose and the 

windward shoulder to the leeside shoulder. 

ACKNOWLEDGEMENTS 

The authors would like to acknowledge Peter 

Gnoffo for his valuable contribution towards the 

sensitivity analyses and assistance with the LAURA 

CFD program. In addition, the authors would like to 

acknowledge Lin Chambers for her assistance with 

the LORAN radiation program. 

Naruhisa Takashima is supported under NASA 

Langley Research Center SAMS contract NAS1-

00135. Ken Sutton is supported under contract 

NAS1-02117 from NASA Langley Research Center 

to National Institute of Aerospace. Michael Wright 

and Dinesh Prabhu were supported under contract 

NAS2-99092 from NASA Ames Research Center to 

ELORET Corp. 

REFERENCES 1

 Lockwood, M, “Titan Aerocapture Systems 

Analysis,” AIAA 2003-4799, July 2003. 2

 Olejniczak, J., et al., “An Analysis of the Radiative 

Heating Environment for Aerocapture at Titan,” 

AIAA 2003-4953, July 2003. 3

 Nelson, H. F., Park, C. and Whiting, E. E., “Titan 

Atmospheric Composition by Hypervelocity 

Shock-Layer Analysis,” Journal of Thermophysics, 

Vol. 5, No. 2, April-June 1991, pp 157-165. 4

 Baillion, M., Pallegoix, J. F. and Soler, J., 

“Huygens Probe Aerothermodynamics,” AIAA 97-

2476, June 1997. 5

 Park, C. and Bershader D., “Determination of the 

Radiative Emission of a Hypersonic Flow 

Simulating the Cassini-Titan Atmospheric Entry 

Probe Environment,” AIAA 90-1558, June 1990. 6

 Baillion, M., “Aerothermodynamic Requirements 

and Design of the Huygens Probe,” Capsule 

Aerothermodynamics, AGARD-R-808, May 1997. 7

 Edquist, K. T., et al., “Aeroheating Environments 

for a Mars Smart Lander,” AIAA 2002-4505, Aug. 

2002. 8

 Justus, C. and Duvall, A., “Engineering- Level 

Model Atmospheres for Titan and Neptune,” 

AIAA-2003-4803, July 2003. 9

 Way, D., Powell, R. and Edquist, K., “Aerocapture 

Simulation and Performance for the Titan Explorer 

Mission,” AIAA 2003 4951, July 2003. 10 Gnoffo, P. A., “Computational 

Aerothermodynamics in Aeroassist Applications,” 

AIAA 2001-2632, 2001. 11 Wright, M. J., Candler, G. v. and Bose, D., “DataParallel Line Relaxation Method for NavierStokes Equations,” AIAA Journal, Vol. 36, No. 9, 

Sept. 1998. 12 Wright, M.J., Loomis, M., and Papadopoulos, P., 

‘Aerothermal Analysis of the Project Fire II 

Afterbody Flow,’ Journal of Thermophysics and 

Heat Transfer, Vol. 17, No. 2, pp. 240-249, 2003. 13 Park, C. “On Convergence of Chemically Reacting 

Flows” AIAA 85-0247, Jan. 1985. 14 Park, C. “A Review of Reaction Rates in High 

Temperature Air” AIAA 89-1740, June 1989. 
82 

15 Chambers, L. H., “Predicting Radiative Heat 

Transfer in Thermochemical Nonequilibrium Flow 

Fields,” NASA TM-4564, Sept. 1994. 16 Whiting, E. E., et al., “NEQAIR96, 

Nonequilibrium and Equilibrium Radiative 

Transport and Spectra Program: User’s Manual,” 

NASA RP-1389, Dec. 1996. 17 Cheatwood, F. M. and Thompson, R., A., “The 

Addition of Algebraic Turbulence Modeling to 

Program Laura,” NASA TM 107758, April 1993. 
83 

AN ANALYSIS OF THE RADIATIVE HEATING ENVIRONMENT FOR AEROCAPTURE AT 

TITAN 

Joseph Olejniczak 

NASA Ames Research Center, Moffett Field, CA 94035 

Michael Wright, Dinesh Prabhu 

ELORET Corp., 690 W. Fremont Ave., Suite 8, Sunnyvale, CA 94087 

Naruhisa Takashima 

AMA Inc., 303 Butler Farm Road, Suite 104A, Hampton, VA 23666 

Brian Hollis, E. Vincent Zoby, 

NASA Langley Research Center, Hampton, VA 23681 

and 

Kenneth Sutton 

National Institute of Aerospace, Hampton, VA 23666 

ABSTRACT 

Details of the radiative heating analysis for the forebody of 

a candidate Titan aerocapture orbiter are presented. The 

radiative heating rates are obtained through a posteriori

analysis of high-fidelity thermochemical nonequilibrium 

flow fields computed using modern techniques of 

computational fluid dynamics. Results from axisymmetric 

and three-dimensional analysis are presented at several 

points on candidate aerocapture trajectories in various 

model atmospheres of Titan. The radiative heating rates are 

found to be up to five times the peak convective heating 

rates, indicating that an accurate knowledge of the 

uncertainty of the radiative heating predictions is critical for 

a Titan aerocapture mission. The results also show that – (1) 

the radiative heating rates are dominated by the violet band 

system of CN, and (2) the gas mixture is optically thin. The 

predicted radiative heating is found to be very sensitive to 

the dissociation rate of molecular nitrogen – a factor of two 

increase in the rate, which is within the experimental 

uncertainty, results in a 25% decrease in the radiative 

heating. 

INTRODUCTION 

one-year system analysis design study1,2 of an 

aerocapture mission to the Saturn moon Titan was 

recently completed. The purpose of the systems study 

was to identify the critical issues for an aerocapture mission, 

not necessarily to provide a complete mission design. The 

purpose of the radiation calculations presented in this paper 

was likewise to assess the impact of radiative heating on the 

system design and identify the further work needed before a 

Titan aerocapture mission could be launched. 

Aerocapture uses a single pass through an atmosphere 

to dissipate excess energy and enter a stable orbit. 

Aerocapture missions offer a mass savings over 

traditional propulsive capture missions if the mass fraction 

of the Thermal Protection System (TPS) needed to protect 

the vehicle during the atmospheric pass is less than the mass 

fraction of the propulsion system needed to slow the vehicle 

and enter orbit. The results of the systems analysis study1

demonstrated that aerocapture missions at Titan provide a 

significant mass benefit compared to propulsive capture 

missions. 

A 70-degree sphere-cone with a 1.875 m maximum 

radius was selected as the baseline design for the orbiter, as 

shown in Fig. 1. The geometry is a scaled version of the 

proposed Mars Smart Lander vehicle.3

 For the nominal 

entry trajectory, the vehicle enters the atmosphere with a 

relative velocity of 6.5 km/s and flies at a constant 16° 

angle-of-attack. For comparable velocities and altitudes at 

Earth, the corresponding radiative heating level would be 

insignificant. 

However, studies of the heating environment of the 

Titan probe Huygens indicated that radiative heating rates 

were 2-3 times greater than convective heating rates.4-6 The 

Huygens probe, part of the Cassini mission to Saturn, is a 

2.70 (m) diameter 60-degree sphere cone that will enter the 

Titan atmosphere on a direct, ballistic trajectory. Since the 

proposed aerocapture orbiter is larger than Huygens and 

flies a longer, lifting trajectory, it was expected that the 

aerocapture orbiter would experience even higher radiative 

heating rates than the Huygens probe. 

Titan has an atmosphere of nitrogen, methane, and 

argon. This gas composition leads to non-intuitive 

thermochemical nonequilibrium effects in the shock layer. 

The methane in the atmosphere dissociates in the 

nonequilibrium shock layer and CN forms. The CN 

molecule is a strong radiator. Therefore, even though the 

A 
84 

X

Z

Y

1.4133 m

1.875 m

0.3281 m

2.0964 m

θc2 = 66.4°

θc1 = 45°

Rs = 0.0914

m

Rn = 0.9124 m

θc = 70°

Afterbody Forebody

Figure 1. Schematic view of the Titan aerocapture orbiter.

entry velocity and post-shock temperatures are relatively 

low compared to Earth entry, there is the potential for 

significant nonequilibrium radiation. 

The gas kinetic and radiation models used in this work 

are based on the work of Nelson et al.,7

 which was 

developed using shock tube data from Park.8 The Park data 

are from only one velocity-pressure condition. A detailed 

collisional-radiative model based on extensive shock tube 

data was developed to design the Huygens probe,4-6 but 

details of the experimental data and the model have not been 

published. 

Because the radiative heating has a first order impact on 

the TPS material selection and mass, even a preliminary 

design of an aerocapture vehicle requires reliable estimates 

of the radiative heating. As a result, high fidelity CFD and 

radiation calculations using the best available 

thermophysical models were made to support the system 

design study. 

This paper discusses these radiative heating predictions 

for the forebody of the Titan aerocapture orbiter. First, a 

description of the methods used to compute the radiative 

heating rates is given. Second, radiative heating results for 

axisymmetric and three-dimensional cases along candidate 

aerocapture trajectories are presented. Finally, the status of 

the ongoing efforts is reported. A series of companion 

papers describes the details and results of the mission 

design.9-16 Takashima et al.17 discusses the convective 

aeroheating environment for the aerocapture vehicle. 

METHODOLOGY 

The flow field and radiation production and transport 

were calculated in an uncoupled manner. In other words, 

Computational Fluid Dynamics (CFD) was used to compute 

the flow field properties used as input to the radiation 

solvers, but the effect of the radiation on the flow field was 

neglected. Previous authors18,19 have used loosely or fullycoupled methods for strongly radiating flows. However, the 

uncoupled approach is appropriate for systems analysis 

trade studies because it provides a conservative estimate of 

the radiative heating at a much lower computational cost 

than a loosely or fully coupled approach, allowing a large 

number of candidate designs to be rapidly analyzed. The 

degree to which the uncoupled approach over-predicts the 

total radiative heating for this application will be discussed 

in a later section. 

The radiation calculations are made by computing the 

radiative flux that reaches a location on the surface of the 

vehicle along a line-of-sight normal to that surface location. 

The calculations required as inputs the species number 

densities, translational temperature, and vibrational 

temperature at each point in the flow field along the line-ofsight. The radiation solvers compute the amount of radiation 

produced and the transport of that radiation along each lineof-sight. 

CFD 

Two different CFD solvers were used to compute the 

flow fields: DPLR19 and LAURA.20 Briefly, both codes 

solve the Navier-Stokes equations including finite rate 

chemistry and vibrational nonequilibrium. DPLR and 

LAURA have been shown to give essentially the same 

results when using identical thermophysical models.17

The chemical reaction rates used were from Nelson et 

al.7

 The rate model in Nelson et al. was developed for 

application to Titan entry and was based on a combination 

of shock tube data from Park8

 and existing reaction rate 

models. Because the amount of nitrogen dissociation has a 

direct impact on the amount of CN formed, some 

calculations were made substituting the newer nitrogen 

dissociation rates from Park22 in order to test the sensitivity 

of the radiative heating to the reaction rate model. The rates 

in Park22 are approximately twice as fast as those in the 

Nelson et al. model. The effect of reaction rate choice on 

radiative heating will be discussed in detail in a later 

section. Most calculations were made using a twelve species 

model (CH4, CH3, CH2, N2, C2, H2, CH, NH, CN, N, C, and 

H). Calculations were also made using an eighteen species 

(CH4, CH3, CH2, N2, N2

+

, C2, H2, CH, NH, CN, CN+, N, N+, 

C, C+

, H, H+, and e) to confirm that ionization effects were 

unimportant. 

RADIATION 

Similarly, two radiation transport codes were used: 

NEQAIR9623 and LORAN.24 Both of these radiation solvers 

calculate the emission and transport of radiation in the flow 

field using the tangent slab approximation. In this 

approximation radiation is assumed to be emitted from 

infinitely thin parallel planes perpendicular to the absorbing 

surface. NEQAIR96 performs the calculations line-by-line, 

computing an absorption coefficient for each atomic line 
85 

Table 1. Methane mass fractions for the three different 

Titan atmosphere models. 

Atmosphere CH4 Mass Fraction 

Minimum 0.0293 

Nominal 0.0173 

Maximum 0.0055 

and molecular rotational line in the vibrational bands. 

LORAN uses a smeared band model with a smoothly 

varying absorption coefficient for the molecular band. 

Comparisons of NEQAIR96 and LORAN results will be 

presented later. 

The tangent slab results computed in either NEQAIR96 or 

LORAN need to be reduced to account for surface curvature 

effects. The exact value of this reduction factor depends on 

the surface geometry and the absorption in the 

 flow field. For a spherical stagnation point in air, values 

from 0.75 to 0.85 have been previously used.25,26 In this 

work, a value of 0.80 was chosen. 

The specific molecular bands considered were N2 1+ 

(B 3Πg–>A 3

Σu

+

), N2 2+ (C 3Πu->B 3Πg), N2 Lyman-BirgeHopfield (a 1

Πg->X 1Σg

+

), CN violet (B 2Σ+->X 2Σ+), CN red 

(A 2

Π->X 2Σ+), H2 Werner (C 1Πu->X 1Σg

+

), and H2 Lyman 

(B 1

Σu

+

->X 1Σg

+

). Also computed was radiation from atomic 

lines of N, C, and H. Updated molecular spectroscopic data 

from Laux27 and atomic spectroscopic data from NIST28

were used in NEQAIR96. 

While CFD calculates the number of molecules and 

atoms in the ground electronic state, radiation is emitted 

from excited electronic states of molecules and atoms. 

Therefore, a collisional-radiative excitation model is needed 

to compute the number of electronically excited molecules 

and atoms and their distribution among vibrational and 

rotational states. 

Since no validated excitation model was available for a 

nitrogen, methane, and argon gas mixture, the excited 

electronic states of the molecules and atoms were assumed 

to be in equilibrium, and therefore populated in a Boltzmann 

distribution governed by the vibrational temperature of the 

gas. While such an assumption is inappropriate for air 

mixtures, it is a reasonable approximation when CN is the 

main radiator emitting in the violet and red bands. Since the 

ground (X 2

Σ+), first excited (A 2Π), and second excited (B 2

Σ+) states of the CN molecule lie close together and are 

distinct with no crossings, it is a reasonable first 

approximation to assume that gas collisional rates will be 

sufficiently high to maintain a Boltzmann distribution 

among the first three electronic states. 

TRAJECTORIES 

The systems analysis group provided four different 

design trajectories, V6.5.2

 The aerothermal environments 

were computed at seven to nine specific points along each 

trajectory. Freestream conditions were taken from an 

atmospheric model8

 developed as part of the Titan mission 

study. Both axisymmetric and three-dimensional CFD 

time (s)

qstag (W/cm2

)

0 250 500 750 0

20

40

60

80

100

120

140

160

Nom. Atm. / Lift-Up - Radiative

Nom. Atm. / Lift-Down - Radiative

Max. Atm. / Lift-Down - Radiative

Min. Atm. / Lift-Up - Radiative

Nom. Atm. / Lift-Up - Convective

Nom. Atm. / Lift-Down - Convective

Max. Atm. / Lift-Down - Convective

Min. Atm. / Lift-Up - Convective

Figure 2. Predicted axisymmetric convective and radiative 

heating rates for four entry trajectories.

calculations were made. Radiative heating values were 

computed at selected points on the body including the 

stagnation point. 

RESULTS 

For aerocapture missions to Titan with entry velocities 

below about 8.5 km/s, our results show that over 90% of the 

radiative heating is from the CN[v] band. The remaining 

radiation is from the CN red band (CN[r]), various N2 

molecular systems, and N atomic lines. Because the 

radiation from CN[v] is relatively short-wavelength, 340–

430 nm, there is a concern that it will penetrate a low 

density TPS material and heat the material in-depth.8

 The 

NEQAIR96 radiation calculations also showed that over the 

entire range of conditions of interest for Titan aerocapture 

the gas is optically thin; i.e. there is no absorption of the 

radiation by the gas in the flow field. 

AXISYMMETRIC CASES

Even though the aerocapture vehicle flow field is threedimensional, initial calculations were made on the orbiter 

geometry assuming an axisymmetric flow field. In this way, 

a large number of trajectory points can be analyzed to 

provide radiative heating sensitivities to the design variables 

being studied by the systems analysts. Figure 2 shows the 

computed axisymmetric stagnation point convective and 

radiative heating rates for four entry trajectories 

corresponding to three different atmosphere models and two 

different flight path angles. 

Since the exact composition of the Titan atmosphere is 

unknown, three different atmospheric models are 

considered: minimum, nominal, and maximum, where the 

nomenclature refers to the atmospheric density at a given 

altitude.10 It is further assumed that the different 

atmospheric models predict differing amounts of methane. 
86 

Table 2. Predicted axisymmetric peak radiative heating rates and radiative heat loads for four entry trajectories. 

Trajectory 

Peak 

Radiative Heating 

(W/cm2

) 

Radiative 

Heat Load 

(J/cm2) 

Peak 

Convective Heating 

(W/cm2) 

Convective 

Heat Load 

(J/cm2) 

Nominal Atm. 

Lift-Up 124 10200 39 5500 

Nominal Atm. 

Lift-Down 84 12090 29 7500 

Maximum Atm. 

Lift-Down 43 8400 28 7700 

Minimum Atm. 

Lift-Up 149 10580 41 5200 

Minimum Atm. 

Lift-Down 100 12480 30 7600

The atmosphere-specific methane mass fractions are listed 

in Table 1. Lift-up indicates a steep entry angle requiring 

full lift to capture to the desired orbit. Lift-down indicates 

a shallow entry angle requiring full lift-down to capture to 

the desired orbit. The lift-up and lift-down trajectories 

thus define the boundaries of the entry corridor. The lift 

of the vehicle is controlled by roll modulation while at 

constant angle-of-attack. 

The corresponding stagnation point heat loads and 

values of peak convective and radiative heating rates are 

given in Table 2. The convective heating rates are below 

50 W/cm2

, however the radiative heating rates are up to 

150 W/cm2 for the minimum atmosphere, lift-up 

trajectory. Peak radiative heating occurs after peak 

convective heating. For the minimum atmosphere, lift-up 

trajectory peak convective heating is at t=253 s and peak 

radiative heating is at approximately t=266 s. The largest 

radiative heat load occurs on the minimum atmosphere, 

lift-down trajectory, and the largest total heat load is for 

the nominal atmosphere, lift-down trajectory. 

Freestream Mole Fraction CH4 (%)

q (W/cm2

)

0 2 4 6 810 0

20

40

60

80

100

120

140

160

Yelle Min (~5%)

Yelle Nominal (~3%)

Yelle Max (~1.5%)

Figure 3. Effect of assumed freestream methane mole 

fraction on axisymmetric stagnation point radiative heating 

predictions for the t = 253 s point of the v = 6.5 km/s entry 

minimum atmosphere, lift-up trajectory.

Interestingly, the trajectory with the peak convective 

heating rate is the trajectory with the peak radiative 

heating rate, yet the trajectory with the maximum 

convective heat load is not the trajectory with the 

maximum radiative heat load. This surprising result is 

because the convective heating depends on the freestream 

enthalpy, while the radiative heating depends on 

temperature and CN concentration distributions through 

the entire shock-layer. 

Although not shown, the predicted radiative heating 

rates are higher along the conical flank of the orbiter than 

at the stagnation point. The peak value is about 230 

W/cm2

. The post-shock temperatures and CN mass 

fractions are similar in the two regions, and the heating is 

higher in the flank region simply because there is a larger 

volume of radiating gas. 

The results shown here are from the DPLR CFD code 

and the NEQAIR96 radiation code using the Nelson et al. 

rate model. The results using the LAURA CFD code and 

the LORAN radiation code with the Nelson et al. rate 

model are within 5% of the DPLR/NEQAIR96 results. 

The radiative heating rates are from CN[v] only. The 

impact of including CN[r], molecular nitrogen bands, and 

atomic nitrogen lines is discussed below. 

Because of chemical nonequilibrium in the shock 

layer, the amount of CN radiation depends strongly on the 

freestream methane concentration. At equilibrium, the 

shock-layer temperatures would be too high for CN to 

exist since the dissociation energy is only 7.65 eV. Figure 

3 shows the calculated radiative heating rates as a 

function of assumed freestream methane concentration for 

the t=253 s point on the minimum atmosphere, lift-up 

trajectory. The freestream velocity is 5761 m/s and the 

density is 1.491e-4 kg/m3

. The amounts of methane 

corresponding to the three atmosphere models discussed 

above are indicated on the figure. The amount of methane 

predicted by the minimum atmosphere model happens to 

result in nearly the peak amount of radiative heating, 

indicating that designing the vehicle using the minimum 

atmosphere is the conservative choice for radiative 

heating. 
87 

Distance from Stagnation Point (m)

CN Number Density (1/m3

)

-0.15 -0.1 -0.05 0 0

2E+20

4E+20

6E+20

8E+20

1E+21

1.2E+21

1.4E+21

1.6E+21

1.8E+21

2E+21

1% CH4

3% CH4

2% CH4

4% CH4

5% CH4

6% CH4

7% CH4

8% CH4

9% CH4

Figure 4. Effect of assumed freestream methane mole 

fraction on post-shock CN number density. 

Distance from Stagnation Point (m)

Temperature (K)

-0.15 -0.1 -0.05 0 0

5000

10000

15000

1% CH4

2% CH4

3% CH4

4% CH4

5% CH4

6% CH4

7% CH4

8% CH4

9% CH4

Figure 5. Effect of assumed freestream methane mole 

fraction on post-shock translational temperature.

Distance from Stagnation Point (m)

Temperature (K)

-0.15 -0.1 -0.05 0

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

10000

11000

T

Tve

T

Tve

Nelson et al. Rates

(qrad = 160 W/cm2

)

Nelson et al. Rates w/ Park N2

(Qrad = 120 W/cm2)

Figure 6. Computed axisymmetric stagnation line temperature 

profiles for two different reaction rate models.

The methane dissociates into its constituent atoms 

immediately downstream of the shock and quickly forms 

CN through various exchange reactions with the 

atmospheric nitrogen. More methane available means 

more CN is produced, as shown in Fig. 4. Therefore, at 

low methane concentrations the amount of radiation is 

proportional to the amount of methane in the atmosphere. 

As the amount of methane increases another effect 

becomes important, namely that the increased amount of 

methane dissociation decreases the post-shock 

temperature (Fig. 5). Even though there are more CN 

molecules, the effective temperature at which the CN 

radiates is lower. The net effect is to decrease the amount 

of CN radiation since the radiation intensity depends 

linearly on the number of radiating molecules but 

exponentially on the vibrational temperature. The 

vibrational temperature is in equilibrium with the 

translational temperature for about 70% of the shock 

layer. 

There is some uncertainty in the thermophysical 

models for the Titan atmosphere. One important rate is 

the nitrogen dissociation rate. Figure 6 shows the effect of 

the value of the nitrogen dissociation rate on the shock 

stand-off distance and the resulting effect on the 

computed radiative heating rate. The two rates differ only 

by a constant factor of about two, and yet the predicted 

radiative heating rates differ by 25%. Additional data are 

needed to reduce this uncertainty. 

THREE-DIMENSIONAL CASES 

Three-dimensional calculations were also made for 

the lift-up, minimum atmosphere trajectory, which has the 

highest radiative heating rates. The distributed radiative 

heating on the forebody of the vehicle is shown in Fig. 7 

for the t=253 s point of the lift-up minimum atmosphere 

trajectory with α=16°. Because the gas is optically thin, 

the radiative heating at the surface is given by the sum of 

250

230

210

190

170

150

130

110

90

qrad (W/cm2)

Leeside

Windside

Figure 7. Computed three-dimensional forebody radiative 

heating rate for the t = 253 s point of the v = 6.5 km/s entry 

minimum atmosphere, lift-up trajectory.
88 

Table 3. Three-dimensional stagnation point radiative heating rates for the t = 253 s point of the v = 6.5 km/s entry 

minimum atmosphere, lift-up trajectory. 

 qrad (CN Violet) 

(W/cm2

) 

CFD 

Code 

# of 

Species Kinetics NEQAIR96 LORAN 

DPLR 12 Nelson et al. 278 246 

DPLR 12 Nelson et al.

& Park N2

214 

DPLR 18 Nelson et al. 289 

LAURA 18 Nelson et al. 

& Park N2

202 175 

the contributions of the individual radiating volumes of 

gas along a line-of-sight perpendicular to the surface. The 

peak radiative heating value is about 260 W/cm2

 is near 

the flow stagnation point where there is the largest 

volume of radiating gas. In the axisymmetric case, the 

largest volume of gas was in the conical flank region and 

the radiative heating rate is about 230 . W/cm2

. The 

axisymmetric stagnation point heating rate of about 150 

W/cm2

 is similar to the value at the center of the spherical 

nose for the three-dimensional calculation. 

Figure 8 shows a comparison of the convective and 

radiative heating rates along the pitch-plane of the 

forebody for the t=253 s point of the lift-up minimum 

atmosphere trajectory. The flow stagnation point is at 

approximately y=0.8 m. The radiative heating is 

significantly larger than the convective heating over the 

entire forebody. 

Table 3 gives the three-dimensional stagnation point 

heating rates for the t=253 point of the lift-up minimum 

atmosphere trajectory for various combinations of CFD 

code, radiation code, and kinetic model. Using 

DPLR/NEQAIR96 the difference between the 12 species 

y (m)

q (W/cm2

)

-1 0 1 0

50

100

150

200

250

Convective

Radiative

Figure 8. Computed pitch-plane forebody convective and 

radiative heating rate for the t = 253 s point of the minimum 

atmosphere, lift-up trajectory.

and 18 species radiative heat transfer is only 4%, 

indicating that ionization effects can be ignored for the 

systems studies. Using the DPLR flow field solution and 

the Nelson et al. model, LORAN predicts approximately 

13% less radiation than NEQAIR96. Considering that 

LORAN uses a lumped band model as opposed to the 

detailed line-by-line method of NEQAIR96, there is 

reasonable agreement between the two codes. For this 

case, the use of the Park nitrogen dissociation rates 

reduces the radiative heating by almost 30%. The results 

from DPLR/NEQAIR96 and LAURA/LORAN are within 

20% when the same kinetics model is used. 

Figure 8 shows that the highest radiative heating 

occurs near the stagnation point of the vehicle. Since at 

higher entry velocities the post-shock temperature in the 

stagnation region will eventually become too high for CN 

to form, it was thought that increasing the entry velocity 

might reduce the amount of radiative heating. Figures 9 

shows the computed radiative heating on the forebody for 

an entry velocity of 10 km/s at the peak heating point. The 

radiative heating levels in the stagnation region are indeed 

reduced from about 250 W/cm2

 to about 160 W/cm2. 

300

275

250

225

200

175

150

125

100

qrad (W/cm2)

Leeside

Windside

Figure 9. Computed three-dimensional forebody radiative 

heating rate for the peak convective heating point (t = 152 s) 

of the v = 10 km/s minimum atmosphere, lift-up trajectory
89 

y (m)

q (W/cm2

)

-1 0 1 0

50

100

150

200

250

Ve = 6.5 km/s

Ve = 10 km/s

Figure 10. Computed pitch-plane forebody radiative heating 

rate for the t = 253 s point of the minimum atmosphere, lift-up 

trajectory and the t = 152 s point of the v = 10 km/s minimum 

atmosphere, lift-up trajectory.

However, for the 10 km/s entry the leeside of the vehicle 

has post-shock conditions similar to the 6.5 km/s 

stagnation point, and the radiative heating rate is as high 

as 220 W/cm2

. A comparison of the radiative heating rates 

for the two entry velocities along the pitch-plane is shown 

in Fig. 10. In fact, the total integrated surface heating rate 

over the orbiter surface increases by 5%. Furthermore, 

this analysis does not include the contribution of the other 

radiators such as N2 and N, which will contribute to the 

radiative heating at this entry velocity. Additionally, the 

convective heating will also be about 3.5 times higher 

compared to the 6.5 km/s entry trajectory. Therefore, 

increasing the entry velocity does not reduce the amount 

of heating to the vehicle. 

 FUTURE WORK 

The prior analyses all assumed that the radiation 

calculation could be separated (decoupled) from the flow 

field calculation. This assumption is appropriate if the 

flow field is adiabatic and the energy lost from radiation 

is small. However, given the high levels of radiative 

heating predicted at relatively low entry velocities, it was 

decided to examine the impact of radiation-fluid coupling 

on the aerothermal heating predictions. 

It is possible to estimate the amount of coupling 

between the radiation and fluid by evaluating the radiative 

loss parameter29, Γ: 

Γ = 2qrad / 1

2 ρ∞V∞

3 (1) 

where qrad is the stagnation point radiative heat flux, ρ∞ is 

the freestream density, and V∞ is the freestream velocity. 

When Γ becomes large (> 1%), the flow is considered to 

be non-adiabatic and coupled to the radiation field. If a 

large amount of flow energy is converted to radiation, the 

shock-layer is cooled and there is a potentially significant 

impact on the fluid dynamics and chemical kinetics of the 

flow. The net effect of the coupling would be to reduce 

the amount of radiative heating. 

The radiative loss parameter for the reference Titan 

aerocapture vehicle was found to exceed 40% near peak 

radiative heating, indicating that the flow is non-adiabatic 

and coupled. By comparison, Γ is about 1% for the Fire-II 

flight experiment, and about 15% for Galileo. Further 

work investigating the effect of coupling is ongoing and 

will be presented in an upcoming paper. Preliminary 

results show that the coupled radiative heating rates are 

still 2-3 times greater than the convective heating rates. 

Preliminary uncoupled afterbody calculations have 

also been made in order to assess the amount of radiative 

heating on the afterbody. The radiative heating analysis is 

still in progress and will be reported in an upcoming 

paper. 

Finally, because the work presented in this paper 

identified that radiative heating is a significant design 

issue for Titan missions, experiments in the NASA Ames 

EAST shock tube are about to begin. The tests will 

provide data over a range of velocity and pressure 

conditions that will be used to either validate the existing 

chemical kinetic and radiation model or construct a new 

collisional-radiative model for application to the Titan 

atmosphere. The development of this new model will 

allow greater confidence in radiative heating predictions 

and quantified uncertainty estimates for Titan entry 

missions. 

SUMMARY AND CONCLUDING REMARKS 

Axisymmetric and three-dimensional thermochemical 

nonequilibrium Navier-Stokes calculations using DPLR 

and LAURA were made for a candidate Titan aerocapture 

orbiter along four design trajectories. The computed flow 

fields were used as inputs to the radiation solvers 

NEQAIR96 and LORAN. The combinations 

DPLR/NEQAIR96 and LAURA/LORAN gave similar 

answers. 

Radiative heating was found to be a significant 

design issue for Titan aerocapture missions. The current 

results are conservative since they do not account for the 

radiative cooling effect, but this work has identified the 

key issues and sensitivities for the radiative heating 

environment. 

The radiation results from the nonequilibrium 

formation of CN in the shock layer around the vehicle 

created in the nitrogen–methane atmosphere. In fact, if the 

flow field were in equilibrium, the radiative heating 

would be negligible compared to the convective heating. 

The radiation field was found to be optically thin for all 

conditions relevant to the aerocapture orbiter. 

It was found that the peak radiative heating rate is 

five times greater than the peak convective heating rate, 

and the radiative heat load is twice as great as the 

convective heat loads for the minimum atmosphere, liftup trajectory. Three-dimensional calculations showed that 

there was significant radiation on the entire forebody. 

The predicted radiative heating was found to be 

sensitive to the choice of the nitrogen dissociation rate. A 
90 

factor of two change in the rates produced a 25% change 

in the predicted heating rate. The factor of two difference 

in the rates is well within the experimental uncertainty of 

the data, indicating more accurate knowledge of the 

nitrogen dissociation rate is needed to reduce the 

uncertainty in the radiative heating predictions. 

This work was performed as part of a preliminary 

system analysis study, and as expected it raised as many 

questions as it answered. Two of these questions, the 

effects of strongly coupled radiation and the impact of 

radiative heating on the afterbody, were mentioned in this 

paper. This work also identified the need for additional 

shock tube data in order to develop the collisionalradiative model necessary for a detailed mission design. 

ACKNOWLEDGMENTS 

M. Wright and D. Prabhu were supported by NASA 

Ames Research Center, through contract NAS2-99092 to 

ELORET. N. Takashima was supported by NASA 

Langley Research Center, through contract to AMA, Inc. 

K. Sutton was supported by NASA Langley Research 

Center, through the National Institute of Aerospace, 

contract number NAS1-02117. 

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Wiley & Sons, 1990. 23 Whiting, E.E., Yen, L., Arnold, J.O., and Paterson, J.A., 

``NEQAIR96, Nonequilibrium and Equilibrium 

Radiative Transport and Spectra Program: User’s 

Manual,” NASA RP-1389, Dec. 1996. 24 Hartung-Chambers, L., “Predicting Radiative Heat 

Transfer in Thermochemical Nonequilibrium Flow 

Fields, Theory and User’s Manual for the LORAN 

code,” NASA TM-4564, Sep. 1990. 25 Reid, R.C., Jr., Rochelle, W.C., and Milhoan, J.D., 

“Radiative Heating to the Apollo Command Module: 

Engineering Predictions and Flight Measurements,” 

NASA Tm X-58091, April 1972. 26 Park, C., “Stagnation-Point Radiation for Apollo 4,” 

AIAA Paper No. 2001-3070, Jun. 2001. 
91 

27 Laux, C.O., “Optical Diagnostics and Radiative 

Emission of Air Plasmas,” HTGL Report No. T-288, 

Stanford University, Aug. 1993. 28 NIST, Washington, DC, “Atomic Spectra Database,” 

(http://physics.nist.gov/cgi-bin/AtData/main_asd), 

2003.

29 Anderson, J.D., Jr., Hypersonic and High Temperature Gas 

Dynamics, McGraw-Hill, 1989.
92 

THERMAL PROTECTION CONCEPTS AND ISSUESFOR AEROCAPTURE AT 

TITAN 

Bernard Laub 

NASA Ames Research Center 

Moffett Field, CA 94035 

ABSTRACT 

A study to develop a conceptual design for an 

aerocapture mission at Titan was conducted by a NASA 

systems analysis team comprised of technical experts 

from several of the NASA centers. Multidisciplinary 

analyses demonstrated that aerocapture could be 

accomplished at Titan with a blunt 70° (half angle) 

rigid aeroshell entering the Titan atmosphere at an 

inertial entry velocity of ≈ 6.5 km/s. Aerothermal 

analyses demonstrated that the peak convective heating 

rates are relatively mild but the radiative heating rates, 

due to shock layer radiation from CN, are significantly 

larger and lie totally in the narrow UV band from 3500 

to 4200 Å. TPS sizing analyses were conducted for a 

broad range of candidate TPS materials and, as 

expected, low density materials are the most attractive 

from a TPS mass standpoint. However, there is 

significant uncertainty associated with the interaction of 

low-density TPS materials with UV radiation (i.e., the 

potential for in-depth absorption). Consequently, the 

preliminary conceptual design adopted a higher mass 

TPS solution until the performance of low-density TPS 

materials can be investigated. 

INTRODUCTION 

Aerocapture is a flight maneuver executed by a 

spacecraft upon arrival at a planet in which atmospheric 

drag is used to decelerate the spacecraft into orbit 

during one atmospheric pass. This contrasts with the 

conventional alternatives of propulsive orbit insertion 

directly into the desired orbit or propulsive insertion 

into a large elliptical orbit followed by a long period of 

aerobraking to reduce the apoapsis altitude. The 

elements of aerocapture to establish a circular science 

orbit are illustrated in Figure 1. 

Although aerocapture has never been attempted in any 

mission to date, it has long been recognized that this 

maneuver can greatly reduce the amount of propellant 

carried by the spacecraft, thereby enabling either larger 

payload mass fractions or smaller launch vehicles from 

Fig. 1. The use of aerocapture to establish a circular 

science orbit about a planetary body1 

Earth. The propellant mass savings of aerocapture 

become especially significant for missions requiring 

large velocity changes for orbit insertion, either because 

a low circular orbit is required or because the approach 

velocity is high. In this context, several missions in 

NASA’s Space Science Strategic Plan are either 

enabled or greatly enhanced by the propellant savings 

afforded by aerocapture technology. These missions 

include Mars Sample Return, Titan Explorer, Neptune 

Orbiter and Venus Surface Sample Return. Other 

severely mass-constrained planetary missions in the 

Discovery Program and any secondary payload 

programs would also clearly benefit from the 

availability of aerocapture technology. 

Through detailed trade studies, the mission analysis 

team determined that aerocapture at Titan could be 

accomplished with a rigid aeroshell configured as a 70 

degree half-angle blunt cone forebody with a lift-todrag (L/D) ratio of 0.25 and a ballistic coefficient 

(M/CDA) of ≈ 90 kg/m2

. The configuration is illustrated 

in Figure 2. Extensive mission analysis studies3 

determined that a viable Titan mission, delivering an 

orbiter with a mass of 590 kg, could be launched 

around December 2010 and, with use of an Earth 

Gravity Assist (EGA) and Solar Electric Propulsion 

(SEP), flight time to Titan could be reduced to 5.9 years 

with an inertial entry velocity of ≈ 6.5 km/s at an 

altitude of 1000 km. 
93 

Fig. 2. Aeroshell configuration for Titan aerocapture2 

TITAN ATMOSPHERE 

The atmosphere around Titan is composed primarily of 

nitrogen with some argon and methane. There is some 

uncertainty about the concentrations of argon and 

methane, which leads to uncertainties in the density 

distribution through the atmosphere. Yelle4

 developed 

engineering models for atmospheric density shown in 

Figure 3. 

Fig. 3 Yelle engineering models for the 

 density of the Titan atmosphere†

The variations in density are attributed to different 

models for methane concentration. The figure indicates 

the range of density variation of the minimum and 

maximum density models relative to the nominal model 

at selected altitudes. The figure also shows the 

minimum altitude range for candidate aerocapture 

trajectories. In these models the methane concentration 

is inversely proportional to mean density. As shown in 

Figure 4, the molar percent CH4 is nearly constant over 

the altitude range where the energy of 6.5 km/s and 10 

 

†

The term “lift down” employed in the figure is 

associated with the overshoot trajectory and indicates 

that the lift vector is in the trajectory plane and is 

always pointing downward during the entire 

aerocapture trajectory. Conversely, “lift up” is 

associated with the undershoot trajectory and indicates 

that the lift vector is always pointing up. 

km/s aerocapture entries would be dissipated. 

Furthermore, the composition does not vary with 

density perturbations. The minimum density 

atmosphere’s methane content drops 50% between the 

surface and 6 km due to CH4 condensation. However, 

CH4 concentrations in the nominal and maximum 

density atmospheres are below the saturation level and 

thus remain constant to the surface. 

AEROCAPTURE TRAJECTORIES 

Starr & Powell5

 evaluated potential aerocapture 

trajectories at Titan with consideration of a range of 

ballistic coefficients and the uncertainties in 

atmospheric density. Guidance, navigation and control 

(GN&C) were limited to controlling the lift vector 

through bank angle modulation. The limiting cases are 

undershoot trajectories, where the lift vector is always 

pointing up, and overshoot trajectories, where the lift 

vector is always pointing down. Four trajectories, 

Fig 4. Models for methane concentration 

in Titan atmosphere 

shown in Figure 5, were selected that bound the limits 

of peak heating rate and maximum total heat load 

CONVECTIVE AND RADIATIVE HEATING 

Convective and radiative heating rates‡ at the stagnation 

point were calculated6

 for the aeroshell configuration 

described previously and the four limiting trajectories 

presented in Figure 5. Stagnation point convective 

heating was calculated using the Fay-Riddell 

correlation7

 and later confirmed with axisymmetric 

Computational Fluid Dynamics (CFD) solutions 

performed with the DPLR code.8

 Non-equilibrium 

radiation calculations were performed with the 

NEQAIR code.9

 As shown in Figure 6, the peak 

 

‡

 Without consideration of the effects of ablation 
94 

Fig. 5. Limiting aerocapture flight trajectories 

stagnation point convective heating rates are less than 

50 W/cm2

 and the undershoot trajectories (lift up) result 

in higher peak heating rates in comparison to the 

overshoot (lift down) trajectories. However, the peak 

stagnation point radiative heating rates are substantially 

larger. For the undershoot trajectories (lift up), peak 

stagnation point radiative heating rates are in the 120-

150 W/cm2

 range. For the overshoot trajectories (lift 

down), peak stagnation point radiative heating rates are 

in the 45-85 W/cm2

 range. 

It is worth noting that the convective heating is 

relatively insensitive to the concentration of methane 

assumed in the atmosphere. However, the radiative 

heating is very sensitive to methane concentration as all 

of the radiative heating is from CN formed in the shock 

layer from the interaction of dissociated methane with 

nitrogen. Consequently, the higher the methane 

concentration in the atmospheric model is (see Fig. 4), 

the higher are the radiative heating rates. While the 

non-equilibrium radiation environment was calculated 

with the most up-to-date chemistry and radiation 

models available, it should be recognized that there is 

significant uncertainty associated with these 

predictions. Shock tunnel tests at NASA Ames are 

planned within the next year to measure CN radiation at 

relevant conditions to enable validation and/or update 

of existing radiation heating models for Titan entry. 

It must also be noted that the overshoot trajectories 

spend a longer time in the atmosphere than the 

undershoot trajectories. Consequently, the total 

convective heat load is typically larger for the 

overshoot trajectories in comparison to the undershoot 

trajectories. That is not necessarily true for the radiative

heating due to the significant differences in the heat 

flux levels predicted for the range of trajectories and

atmospheric models considered. This is demonstrated in 

Table 1, which summarizes the total convective and 

0

20

40

60

80

100

120

140

160

0 100 200 300 400 500 600 700 800

Nom. atm / Lift-Up / Convective

Nom. atm / Lift-Up / Radiative

Nom. atm / Lift-Down / Convective

Nom. atm / Lift-Down / Radiative

Max. atm / Lift-Down / Convective

Max. atm / Lift-Down / Radiative

Min. atm / Lift-Up / Convective

Min. atm / Lift-Up / Radiative

Stagnation Point Heat Flux (W/cm2

)

Time (s)

Fig. 6. Stag point heating for limiting trajectories 

radiative heat loads for the four limiting trajectories 

considered. Note the similarities in the convective heat 

loads for the two overshoot and two undershoot 

trajectories. However, the radiative heat loads exhibit a 

dependence on methane concentration, that is, the 

minimum density atmosphere model has the highest 

methane concentration (see Fig. 4) while the maximum 

density atmosphere model has the lowest methane 

concentration. Table 1 suggests that total radiative heat 

load is dependent on time in the atmosphere (for 

example, compare the values for the lift up and lift 

down trajectories for the nominal density atmosphere) 

but a stronger dependence on the atmospheric model 

(e.g., compare the values for the maximum atmosphere 

on a lift up (undershoot) trajectory with the values for 

the minimum atmosphere on a lift down (overshoot) 

trajectory). 

Table 1. Total heat loads for limiting aerocapture 

trajectories 

Atmosphere model/ 

aerocapture trajectory 

Convective 

heat load 

(J/cm2

) 

Radiative 

heat load 

(J/cm2) 

Nominal atm/Lift up 5,500 10,021 

Nominal atm/Lift down 7,500 12,090 

Maximum atm/Lift 

down 

7,700 8,393 

Minimum atm/Lift up 5,200 15,769 

CANDIDATE TPS MATERIALS 

Given the range of convective and radiative heating 

described above, a range of thermal protection materials 

was identified as candidates for Titan aerocapture 

applications. The candidate TPS materials are 

summarized in Table 2. 
95 

Table 2. Candidate forebody TPS materials for Titan aerocapture§

The list encompasses well-characterized materials such 

as the Shuttle tiles, SLA-561V (used on Mars Viking,

Mars Pathfinder, and Mars Exploration Rover) and 

fully dense carbon phenolic (used on the Pioneer Venus 

and Galileo entry probes). It also includes some more 

recent TPS materials such as PICA (used as the 

forebody TPS on Stardust), SIRCA (used as the 

Backshell Interface Plate on Mars Pathfinder and Mars 

Exploration Rover), and the Genesis Concept 

(employed by Lockheed-Martin as the forebody TPS on 

§ The list of candidate materials was assembled based 

on predictions of convective heating for Titan 

aerocapture but prior to the availability of predicted 

radiative heating. The magnitude of the predicted peak 

radiative heating rates would eliminate some of the 

candidate materials from serious consideration. 

the Genesis spacecraft). The list has some 

developmental materials such as low-density silicone 

and phenolic-based composites being developed by 

Applied Research Associates (ARA), TUFROC, a 

multilayer, non-ablative concept being developed by 

NASA Ames Research Center, and Acusil I, a moderate 

density silicone in honeycomb developed by ITT 

Aerotherm. 

Most of these materials are organic resin-based 

composites that will pyrolyze when heated leaving a 

carbonaceous char at the surface. Table 2 provides the 

virgin density of these materials, a general description 

of their composition and construction, important optical 

properties (where known or can be estimated), some 

comments on their performance limits (where known or 

can be estimated), and some comments on performance 

uncertainties as perceived by the author. 

Material Shuttle tiles SLA-561V SRAM 14 (ARA) SRAM 17 (ARA) 

 

Density (g/cm3) 0.192-0.352 0.256 0.224 0.272 

Description Low-density, glassbased ceramic tile 

with glass-based 

coating. 

Low-density cork 

silicone composite in 

Flexcore 

honeycomb. 

Low-density cork 

silicone composite 

fabricated w/strip 

collar bonding 

technique. 

Low-density cork 

silicone composite 

fabricated w/strip 

collar bonding 

technique. 

 

Optical properties 

 Solar absorptance - 0.50 - 0.50 - 0.50 

 Total hemis. 

emittance 

 - 0.78 - 0.78 - 0.78 

 

Performance limits 44 W/cm2

 (certified) Ablative. 

Char spall at p > 

0.25 atm. 

No (little) recession 

at 

q& < 100 W/cm2. 

Ablative. 

No (little) recession 

at 

q& < 100 W/cm2. 

Differential 

recession between 

composite ad 

interface strips may 

cause boundary layer 

transition. Other 

limits (if any) 

unknown. 

Ablative 

No (little) recession 

at 

q& < 100 W/cm2

Differential 

recession between 

composite ad 

interface strips may 

cause boundary layer 

transition. Other 

limits (if any) 

unknown. 

 

Uncertainties May not be adequate 

for current estimates 

of peak convective 

heating. Will 

probably be 

transparent to CN 

radiation; a showstopper 

Low density, porous 

material. In-depth 

radiant transmission 

may be important. 

Low density, porous 

material. In-depth 

radiant transmission 

may be important. 

Low density, porous 

material. In-depth 

radiant transmission 

may be important. 
96 

Table 2. Candidate forebody TPS materials for Titan aerocapture (continued) 

From the TPS standpoint, total heat loads for 

aerocapture are typically much larger than for direct 

entry. Consequently, the lowest mass TPS solutions 

will be good insulators, a characteristic that is usually 

associated with low-density materials. However 

insulation performance must be balanced with ablation 

performance, for example, too much surface recession 

can lead to alterations in shape that can affect 

aerodynamic performance. High-density materials are 

usually employed to minimize surface recession. 

However, the peak heating rates predicted for Titan 

aerocapture are within the capabilities of several of the 

low-density candidate materials, that is, they can 

provide good insulation performance with minimal 

surface recession. Consequently, they are considered 

the most attractive candidate materials from the 

standpoint of minimizing TPS mass. 

TPS MASS ESTIMATES 

To provide an estimate of TPS mass, preliminary 

forebody TPS sizing analyses were done for the four 

limiting aerocapture trajectories shown in Figure 5 for 

many (but not all) of the candidate materials listed in 

Table 2. The analyses were limited to the stagnation 

point heating shown in Figure 6 and, for purposes of 

estimating TPS mass, it was assumed that the nominal 

(without margin) stagnation point TPS thickness is

Material SIRCA PICA TUFROC SRAM 20 (ARA) 

 

Density (g/cm3) 0.192-0.352 0.256 Varies with layer 

sizing 

0.32 

 

Description Low-density ceramic 

tile impregnated with 

silicone resin. 

Low-density carbon 

fiberform partially 

filled with phenolic 

resin. 

Multilayer composite: 

carbon 

fiberform/AETB tile 

with high emissivity, 

high temperature 

surface treatment 

Low-density cork 

silicone composite 

fabricated w/strip 

collar bonding 

technique. 

 

Optical properties 

 Solar absorptance - 0.28 - 0.90 - 0.90 - 0.50 

 Total hemis. 

emittance 

- 0.92 - 0.80 - 0.90 - 0.78 

 

Performance limits Ablative. 

No (little) recession 

at q& < 100 W/cm2

. 

Rapid recession 

(melting) at higher 

heat fluxes. 

Fabricated as tiles. 

Ablative. 

No recession in nonoxidizing atmosphere 

at 

q& < 1000 W/cm2. 

Excellent low 

density ablator, but 

not best insulator. 

Non-ablative. 

May be usable to 

q& ~ 300 W/cm2. 

Has to be fabricated 

as tiles. 

Ablative. 

No (little) recession 

at 

q& < 100 W/cm2. 

Differential recession 

between composite 

ad interface strips 

may cause boundary 

layer transition. 

Other limits (if any) 

unknown. 

 

Uncertainties Low-density, glassbased material. 

In-depth radiant 

transmission may be 

important. 

Low density, porous 

material. In-depth 

radiant transmission 

may be important. 

Heat flux limit 

currently uncertain. 

Spectral emittance 

data on coating shows 

very high emittance 

at CN violet 

wavelengths. Will 

absorb (not transmit) 

UV radiation. 

Low density, porous 

material. In-depth 

radiant transmission 

may be important. 
97 

Table 2. Candidate forebody TPS materials for Titan aerocapture (concluded) 

applied uniformly on the forebody. It was also assumed 

that the TPS is adhesively bonded to a rigid 

substructure consisting of 0.0376 mm thick graphite 

polyimide facesheets (front and back) on a 31.75 mm 

thick aluminum honeycomb. The density for the 

honeycomb and graphite polyimide facesheets is 0.069 

g/cm3

 and 1.0 g/cm3, respectively. It was assumed that 

all materials are at a uniform temperature of –74.8°C at 

atmospheric interface. It was also assumed that all 

candidate materials absorb CN radiation at the surface 

and perform as thermochemical ablators, that is, no 

spall or melt runoff. Analyses were performed to 

determine the thickness required for each candidate 

material to limit the maximum bondline temperature to 

250°C. It should be noted that different individuals did 

the analyses for different materials, but all used 

comparable analysis tools that address the fundamental 

physical and chemical mechanisms associated with the 

thermal/ablation performance of these materials in the 

Titan atmosphere. Some of the materials models are 

very mature and have been validated with extensive 

laboratory and arc jet test data. In contrast, some of the 

other materials are relatively new and their models are 

based on limited laboratory and arc jet test data. 

The results of the analyses demonstrated that, for the 

four trajectories considered, the TPS thickness is 

significantly larger for the overshoot trajectories (lift 

down) than for the undershoot trajectories (lift up). This 

was the case for all of the candidate materials 

considered. Furthermore, maximum bondline 

temperature is attained during heat soak, that is, after 

the end of aerodynamic heating. 

Material PhenCarb-20 

(ARA) 

Acusil I Carbon phenolic Genesis concept 

 

Density (g/cm3) 0.32 0.48 1.45 Varies with layer 

sizing 

Description Low-moderate 

density phenolic 

composite 

fabricated w/strip 

collar bonding 

technique. 

Moderate density 

filled silicone in 

Flexcore 

honeycomb. 

Fully dense tapewrapped or chopped 

molded heritage 

material. 

Carbon-carbon 

facesheet over 

carbon fiberform 

insulator. 

Optical properties 

 Solar absorptance - 0.50 0.54 - 0.90 - 0.90 

 Total hemis. 

emittance 

- 0.78 0.95 - 0.90 - 0.90 

Performance limits Ablative. 

No recession in 

non-oxidizing 

environment. 

Differential 

recession between 

composite ad 

interface strips may 

cause boundary 

layer transition. 

Other limits (if any) 

unknown. 

Ablative. 

No recession (in 

air) at 

q& < 100 W/cm2. 

Will recede in air at 

higher heat fluxes 

(assumed by 

oxidation). 

Ablative. 

No recession 

expected for Titan 

aerocapture 

environment. 

Ablative. 

No recession 

expected for Titan 

aerocapture 

environment. 

Uncertainties Higher density and 

higher char yield of 

phenolics may 

mitigate in-depth 

radiant absorption 

at the penalty of 

higher thermal 

conductivity. 

Higher density and 

good char yield 

may mitigate indepth radiant 

absorption. 

Not as good an 

insulator as lower 

density materials. 

Difficult to 

fabricate in 

thicknesses > 2 

inches. 

High density and 

high char yield will 

mitigate in-depth 

radiant absorption. 

Not a very good 

insulator. Without 

recession, will be 

very heavy (a poor 

choice for this 

application). 

High-density carbon 

facesheet will 

mitigate in-depth 

radiant absorption. 

Carbon fiberform 

not the best 

insulator, but better 

than fully dense 

carbon phenolic. 
98 

Table 3. Preliminary forebody TPS sizing for Titan aerocapture 

Maximum atmosphere - Lift Down 

Convective Heat Load = 7,700 J/cm2 

Radiative Heat Load = 8,393 J/cm2

Nominal atmosphere – Lift Down 

Convective Heat Load = 7,500 J/cm2

Radiative Heat Load = 12,090 J/cm2

Candidate 

TPS Material Thickness 

(cm) 

Areal weight 

(g/cm2

) Thickness 

(cm) 

Areal weight 

(g/cm2

) 

SLA-561V 2.44 0.626 2.43 0.622 

SRAM 14 1.57 0.353 1.55 0.348 

SRAM 17 1.93 0.526 1.93 0.526 

SRAM 20 2.08 0.667 2.08 0.667 

PhenCarb-20 2.29 0.696 2.34 0.711 

TUFROC 4.88 1.117 5.13 1.181 

Genesis --- --- 5.51 1.298 

PICA 5.94 1.591 5.82 1.557 

Carbon phenolic 8.70 13.084 8.76 13.167 

The results of the TPS sizing analyses are summarized 

in Table 3 which only shows the thickness and areal 

weight requirements for the overshoot trajectories 

because, as stated previously, they exceeded the 

requirements for the undershoot trajectories. 

Table 3 illustrates that the thickness and areal weight 

for any candidate TPS material is similar for the two 

overshoot trajectories despite the significant difference 

in the total radiative heat load. Examination of Figure 6 

will illustrate that the predicted radiative heating for the 

nominal atmosphere model results in higher peak heat 

fluxes in comparison to the maximum atmosphere 

model, but for a shorter period of time. As it was 

assumed that all of these materials are surface 

absorbers, the higher heat fluxes result in higher surface 

temperatures. As surface re-radiation is the 

predominant energy accommodation mechanism for all 

of these candidate materials, they are all more 

“efficient” at higher heat fluxes as long as they do not 

experience significant surface recession. 

Table 3 also illustrates that, in general, TPS areal 

weight increases with material density. This is clearly 

evident by comparing the areal weight requirements 

between the candidates in the SRAM family where the 

composition and construction are similar but the virgin 

densities are different. Also, note that the areal weight 

of PhenCarb 20 is similar to SRAM 20. However, that 

is coincidental; the materials have the same virgin 

density but different compositions, as SRAM uses a 

silicone resin binder and PhenCarb, a phenolic resin 

binder. The thermal conductivity of the PhenCarb 

family should be higher than that of the SRAM family, 

but that is compensated for by the smaller surface 

recession experienced by the PhenCarb in comparison 

to the SRAM. 

This tradeoff is clearly demonstrated when one 

compares the areal weight of PICA with SRAM17 

where the virgin densities are similar. PICA, composed 

primarily of carbon, has a much higher thermal 

conductivity than SRAM17 and this is reflected in the 

areal weights, with the areal weight of PICA 

approximately three times that of SRAM17. The fully 

dense carbon phenolic result clearly makes the point 

because, from the standpoint of composition, it is 

similar to PICA. Yet the almost six-fold increase in 

virgin density is accompanied by a significant increase 

in thermal conductivity, which is reflected in the areal 

weight requirements. The key point is that the materials 

that are primarily carbonaceous are excellent “ablators” 

but not good insulators. They are best used in 

environments with very high heating rates because they 

are most efficient when they ablate. The heating rates 

for Titan aerocapture trajectories are not sufficiently 

severe to allow these materials to perform efficiently. 

The low density, low thermal conductivity materials are 

better choices. 

The Genesis and TUFROC candidates need to be 

viewed differently. The Genesis TPS employs a thin 

carbon-carbon facesheet over a carbon fiberform 

insulator. The carbon-carbon is robust and will not 

experience any recession for the range of Titan 

aerocapture heating environments. Although the carbon 

fiberform is not the best insulator, it is one of the few 

materials that can be employed in contact with the 

carbon-carbon because such a material must be stable at 

the surface temperatures that the C-C facesheet will 

attain. The TUFROC concept is different in some 
99 

aspects. It is comprised of a carbon fiberform insulator 

bonded to an AETB low-density ceramic insulator. The 

two-layer composite has a high temperature, high 

emissivity, low catalytic efficiency surface treatment. 

The assembly is intended to be non-ablative, but the 

upper limit of the surface treatment is yet to be 

demonstrated. The sizing of TUFROC, a two-step 

process as defined by the carbon fiberform thickness, is 

governed by the upper temperature limit of the AETB 

ceramic. Once that is defined, the AETB thickness is 

sized to limit bondline temperature. 

As a consequence of the carbon fiberform minimum 

thickness requirement for TUFROC, the areal weight of 

this concept is larger than for most of the other 

candidate materials. Existing data suggest that the 

coating can handle the range of heat fluxes anticipated 

for the range of Titan aerocapture trajectories 

considered. Furthermore, existing optical properties 

data on the coating indicate that it will absorb UV 

radiation. The results for the Genesis concept indicate 

that its areal weight will be slightly larger than that for 

TUFROC, primarily because the carbon fiberform is 

not as good an insulator as a composite of carbon 

fiberform over a low-density AETB ceramic. 

TPS PERFORMANCE UNCERTAINTIES 

Of some concern is the interaction of CN radiation with 

low density, porous TPS materials. Figure 7 illustrates 

the spectral distribution of the predicted¶

 CN radiation 

where it is seen that almost all the radiation lies in a 

relatively narrow band in the ultraviolet (UV) with the 

peak at ≈ 3800Å (0.38μm). Studies conducted during 

the 1980s, evaluated the performance of dozens of 

ablative materials exposed to high-energy lasers. The 

types of materials evaluated spanned the range from 

low-density organic resin composites to fully dense 

carbon-carbon composites. Materials were tested with 

both continuous wave (CW) and repetitively pulsed 

(RP) lasers at wavelengths from the visible (0.53μm) to 

the infrared (10.6μm). While material performance was 

strongly dependent on the type of material and the 

irradiance (heat flux) it was exposed to, the data also 

suggested a general trend where material performance 

degraded at the shorter wavelengths. Further studies 

demonstrated that the materials did not become semitransparent at the shorter wavelengths, but rather the 

absorption length became larger as the wavelength got 

shorter. It should be noted that none of the materials 

that are TPS candidates for Titan aerocapture were 

evaluated under these laser studies. 

 

¶

 Nonequilibrium NEQAIR calculations 

Based on the trend observed in the laser studies, the 

performance of the most attractive TPS candidates for 

Titan aerocapture is, at best, uncertain. If these lowdensity materials have significant absorption lengths to 

radiation at UV wavelengths, there is the potential that 

the materials may spall (blow off the surface layer due 

to in-depth pyrolysis). It can be argued that when these 

materials pyrolyze they will form a carbonaceous 

0

200

400

600

800

1000

1200

0.30 0.35 0.40 0.45 0.50

Steradiance (W/cm2

-μm-ster) x 102

Wavelength (μm)

Time = 266 sec

Fig. 7 Spectral distribution of CN radiation 

surface char with different absorption characteristics 

than the virgin material; the char will be a surface 

absorber and/or scatter the incident radiation. That is 

certainly possible and, in fact, anticipated if the char 

has sufficient time to form and stays attached. That is 

the major issue because it is possible that in-depth 

absorption leading to periodic spall would continually 

remove whatever surface char is in the process of being 

formed. 

It is also possible that pyrolysis gases injected into the 

boundary layer could scatter the incoming CN radiation 

or absorb it and re-radiate at longer wavelengths. 

Again, that is possible but it is very difficult to 

demonstrate, even analytically, due to uncertainties in 

the composition of the pyrolysis products and their 

absorption characteristics. 

The only way to evaluate the performance of these 

materials when exposed to UV radiation is through 

experiments. Unfortunately, a good experimental 

simulation is not simple. The ideal experiment would 

expose these candidate materials to a radiation source at 

the wavelengths of interest and at relevant heat fluxes. 

All of the high-energy CW lasers operate in the 

infrared. There are excimer lasers that operate at the 

wavelengths of interest, but they are pulsed lasers that 

produce extremely high heat fluxes over a very short 

duration. There are some CW Argon-ion and Krypton 

lasers that produce radiation in the range of interest, but 

they are very low power. 

It has been suggested that the materials be tested with a 

solar source. Examination of the black body spectral 
100 

distribution of solar radiation illustrates that only a 

small fraction (<15%) of the energy lies in the UV band 

of interest with a significant fraction in the IR. 

Furthermore, the atmosphere filters out much of the UV 

radiation. Testing with a solar source would expose the 

materials to broadband radiation and the potential for 

in-depth absorption of UV radiation could be masked 

by char formation promoted by surface absorption of IR 

radiation. 

Tests could also be done with a high power xenon lamp 

that simulates a 5800 K source. Examination of the 

black body spectral distribution from such a source 

illustrates that only a small fraction (≈15%) lies in the 

band of interest. It is possible to filter the radiation from 

this lamp to allow only the UV radiation to reach the 

material sample, but at the sacrifice of much of the 

power. 

The best option is the use of a commercially available 

Mercury-Xenon Lamp, which will allow the testing of 

2.54 cm (1 inch) square specimens at the required 

energy flux levels. Mercury-Xenon short arc lamps 

are dosed with an exact amount of mercury and xenon 

gas. Due to the high operating pressure of mercuryxenon, the spectrum shows only traces from the broad, 

visible and infrared dominated xenon gas spectrum. The 

four main mercury lines dominate the spectrum, as 

shown in Figure 8. The first three lines provide an 

excellent simulation of the predicted CN radiation at 

Titan. The visible lines between 500 and 600 nm are 

easily filtered out, and can be included or not as 

desired. A simple cold filter can eliminate the visible 

and infrared lines. 

A more fundamental approach would measure the 

relevant optical properties of the virgin (and char) 

candidate materials using a spectrophotometer and an 

integrating sphere. Such data would be very valuable 

but would require making (some) measurements on 

very thin samples. Slicing low-density materials into 

thin sections (≈ 1 mm) would be a challenge. 

From the above discussion it should be apparent that 

validating adequate performance of the best TPS 

candidates for Titan aerocapture would not be simple 

due to experimental limitations. Some combination of 

the aforementioned tests must be done and the resultant 

data will need to be carefully evaluated before one can 

design a TPS for Titan aerocapture with confidence. If 

adequate performance of the most attractive candidate 

materials when exposed to CN radiation cannot be 

Fig. 8. Spectral output of Mercury-Xenon lamp shows 

good simulation of CN lines. 

experimentally demonstrated, any TPS design 

employing these materials will have to include 

significant “margin” to mitigate the risk. 

PRELIMINARY TPS CONCEPTUAL DESIGN 

The charter of the NASA aerocapture systems analysis 

team is focused on identifying and prioritizing the most 

critical technology gaps to enable aerocapture to be 

utilized for Solar System Exploration (SSE) missions. 

The benefits of employing aerocapture for Titan 

Explorer can only be quantified through comparison 

with a design that employs chemical propulsion for 

orbit insertion. To enable that comparison, the 

aerocapture systems analysis team developed a 

preliminary conceptual design of an aerocapture 

“system” for Titan Explorer so that comparisons could 

be made of launch mass, payload mass, and system 

complexity. An estimate of TPS mass was required to 

support this conceptual design. For this purpose the 

systems analysis team preliminarily selected TUFROC 

for the forebody TPS. As seen in Table 3, the nominal 

areal weight of TUFROC (only sized at the stagnation 

point) is ≈ 1.18 g/cm2

 for the worst-case trajectory 

(nominal atmosphere, lift down). The surface area of 

the forebody aeroshell is ≈ 12.59 m2

, which results in a 

nominal forebody TPS mass of ≈ 149 kg. The TPS 

sizing studies identified several candidate materials 

that, potentially, would provide lower TPS mass 

solutions, in some cases by factors of 2-3. However, all 

of these attractive candidate materials need to be 

experimentally evaluated to ensure that potential indepth radiant absorption will not significantly degrade 

material performance. 
101 

SUMMARY 

A NASA systems analysis team developed a rigid 

aeroshell conceptual design for orbit insertion around 

Titan using aerocapture. Candidate TPS materials were 

identified and preliminary TPS sizing studies were 

conducted using predicted stagnation point convective 

and radiative heating for limiting aerocapture 

trajectories. The results of these analyses demonstrated 

that existing low-density organic resin composites 

provide low TPS mass solutions for this mission. 

However, the potential for in-depth absorption of CN 

radiation can degrade the performance of these 

materials. The interaction of these materials with UV 

radiation needs to be experimentally evaluated, but the 

options for conducting the appropriate ground test 

experiments simulating UV radiation at relevant heat 

fluxes are limited. Therefore, the preliminary 

conceptual design incorporated a TPS concept 

(TUFROC) that has been demonstrated to absorb UV 

radiation at the surface, but requires a significantly 

larger TPS mass. 

ACKNOWLEDGEMENTS 

The author wishes to express his appreciation to Dr. Y. 

-K. Chen (NASA Ames Research Center) who 

conducted the TPS sizing analyses for SLA-561V, 

PICA, and TUFROC, to Dr. Gary A. Allen 

(ELORET/NASA Ames Research Center) who 

conducted the TPS sizing analyses for fully-dense 

carbon phenolic, to Mr. William M. Congdon (Applied 

Research Associates) who conducted the TPS sizing 

analyses for SRAM14, SRAM17, SRAM20 and 

PhenCarb20, and to Mr. John A. Dec (NASA Langley 

Research Center) who conducted the TPS sizing 

analyses for the Genesis TPS concept. The author also 

wishes to acknowledge Dr. Donald M. Curry (NASA 

Johnson Space Center) and Dr. F. Neil Cheatwood 

(NASA Langley Research Center) who were consulted 

in assembling the list of candidate TPS materials for 

Titan aerocapture. The author also wishes to 

acknowledge the special contributions of Dr. Susan M. 

White (NASA Ames Research Center) who researched 

available devices to simulate the Titan radiation 

environment and identified the Mercury-Xenon lamp. 

REFERENCES 

1. Lockwood, M.K. “Titan Aerocapture Systems 

Analysis Overview,” NASA Langley Research 

Center, Aerocapture Systems Analysis Review, Jet 

Propulsion Laboratory, Pasadena, California. 

August 29-30, 2002. 

2. Hrinda, G.A., “Aeroshell Structure,” NASA 

Langley Research Center, Aerocapture Systems 

Analysis Review, Jet Propulsion Laboratory, 

Pasadena, California. August 29-30, 2002. 

3. Noca, M., “Mission Analysis,” Jet Propulsion 

Laboratory, Aerocapture Systems Analysis 

Review, Jet Propulsion Laboratory, Pasadena, 

California. August 29-30, 2002. 

4. Yelle et al.: “Engineering Models for Titan’s 

Atmosphere”, in Huygens Science, Payload and 

Mission, 

ESA SP-1177, 1997 

5. Starr, B.R. and Powell, R., “Titan/Neptune 

Aerocapture System Study,” NASA Langley 

Research Center, Hampton, VA, March 29, 2002.

6. Olejniczak, J. and Allen, G.A., Jr., “Radiative 

Heating/TPS Sizing Results,” NASA Ames 

Research Center, May 2, 2002.

7. Fay, J.A. and F.R. Riddell, “Theory of Stagnation 

Point Heat Transfer in Dissociated Air,” J. Aero. 

Sci., 25, (2), 73-85, 1958.

8. Wright, M.J., G.V. Candler, and D. Bose, "DataParallel Line Relaxation Method for the NavierStokes Equations," AIAA Journal, Vol. 36, No. 9, 

pp. 1603-1609, Sep. 1998.

9. Whiting, E.E., Park, C., Liu, Y., Arnold, J.O., and 

Paterson, J.A., "NEQAIR96, Nonequilibrium and 

Equilibrium Radiative Transport and Spectra 

Program: User's Manual," NASA RP-1389, Dec. 

1996
 

102

STRUCTURAL DESIGN OF THE TITAN 

AEROCAPTURE MISSION 

Glenn A. Hrinda 

NASA Langley Research Center, Hampton VA

ABSTRACT 

 A major goal of NASA’s planetary exploration 

efforts is to create affordable spacecraft capable of 

delivering science experiments for long duration 

periods. To help achieve this goal the aerocapture 

technique for slowing a spacecraft has been 

investigated and appears to produce less vehicle mass 

then an all-propulsive mission. A conceptual 

spacecraft was designed and studied for an 

aerocapture mission to Titan, Saturn's largest moon. 

The spacecraft is an Orbiter/Lander combination that 

separates prior to aerocapture at Titan. The structural 

challenges faced in the design will be discussed as 

well as optimization sizing techniques used in the 

Orbiter’s aeroshell structure. Design trades required 

to optimize the structural mass will be presented. 

Member sizes, concepts and material selections will 

be presented with descriptions of load cases and 

spacecraft structural configurations. Areas of 

concern will be highlighted for further investigation. 

This study involved the colaberation efforts of NASA 

representatives from Langley Research Center 

(LARC), Jet Propulsion Lab (JPL) and Ames 

Research Center (ARC). The concept design 

borrowed from existing flight hardware as much as 

possible. 

INTRODUCTION 

 The structural sizing for a conceptual aerocapture 

spacecraft to Titan was required to obtain mass 

estimates based on current sizing methods. Finite 

element analysis (FEA) and HyperSizerTM sizing 

software was used to model the launch stack 

assemble that included the Propulsion Module (PM), 

the Orbiter and the Lander. The Orbiter spacecraft 

performs aerocapture at Titan and is designed to 

withstand atmospheric heating. The Lander is a 

sphere-cone and was considered as a concentrated 

mass. No aeroshell design analysis was performed 

on the Lander. The launch vehicle used was a 

Boeing Delta IV heavy with a 4 meter fairing. The 

spacecraft integration into a Delta IV heavy launch 

vehicle was achieved through trade studies focusing 

on mission performance necessary for an aerocapture 

mission. Three primary design objectives were: 

minimum structural mass, dynamic modes at launch 

were meet or exceeded and stress levels were within 

margin with minimal deflections. Load cases and 

frequency minimums at launch came from the Boeing 

Payload Planners Guide. Maximum loading at 

launch and during entry at Titan was used to design 

the spacecraft structure. The lowest predicted natural 

dynamic modes were investigated to identify any low 

frequency problems with the spacecraft. 

 The structural design used composites for the 

Orbiter aeroshell and a truss system to join the stack 

components. Modeling efforts were kept as simple as 

possible to shorten modifications occurring as the 

design progressed. HyperSizerTM sizing software 

was found beneficial in sizing the Orbiter’s aeroshell. 

The software’s ability to optimize composite sections 

without refining mesh densities and geometry was 

demonstrated throughout the design’s progress. 

NOMENCLATURE/ABBREVIATIONS 

CG Center of Gravity 

FEA 

FEM 

Finite Element Analysis 

Finite Element Model 

HGA High Gain Antenna 

NSM Non-structural mass 

PM Propulsion Module 

TPS Thermal Protection System 

CONCEPTUAL TITAN DESIGN 

 The Titan aerocapture spacecraft is a stacked 

configuration requiring a three component stack 

consisting of a Propulsion Module (PM), an Orbiter 

and Lander. The three separate spacecraft were 

combined to form a launch stack capable of fitting 

into a 4 meter Delta IV fairing. Each vehicle must be 

able to separate during the Titan aerocapture mission 

sequence. The PM provides all thrust maneuvering 
 

103

to get the Orbiter and Lander near Titan. The Orbiter 

and Lander will then separate from the PM and then 

from each other. The Lander descends to the Titan 

surface in a Huygens type aeroshell. The Orbiter 

continues through the thin Titan atmosphere and 

begins aerocapture until achieving its mission orbit. 

The Orbiter was the only vehicle designed to take 

advantage of aerocapture. Once in orbit around 

Titan, the Orbiter will support an on orbit relay 

station for the Lander. 

 The conceptual spacecrafts were used as a 

baseline to test design and analysis methods used 

among the various NASA centers involved with 

aerocapture vehicle designs. This study focused on 

the weight reduction and strength requirements of the 

major load carrying structural members. The design 

attempts to maintain an axial load path direction 

starting with the Lander, into the Orbiter through its 

payload pallet and heatshield and final into the PM. 

 

 

Figure 1. Launch Stack 

 The structural analysis performed in this study 

helped verify the stack arrangement and size the 

Orbiter aeroshell and support structure for the generic 

Lander. Investigating various stack arrangements 

showed that a truss would provide the lightest 

structure for supporting the Lander. A truss was also 

used for the PM to Orbiter adapter structure. The 

study used a launch load envelop for the Delta IV 

heavy. The Orbiter maximum diameter was set to 

3.75 meters and used a heatshield cone and biconic 

backshell as shown in figure 1. The PM was 

modeled to include its stiffness contribution in 

determining overall stack frequency during launch. 

Launch loads were taken from the Boeing Delta IV 

Payload Designers manual (ref. 1). 

STACK CONFIGURATIONS 

 The stacking sequence of the PM, Orbiter and 

Lander was decided upon after several trials of the 

three components arranged in different 

configurations. Each configuration had its abilities 

compared with each other until the stack shown in 

figure 1 was chosen. This arrangement was used 

after various stack sequences were attempted to find 

a stack able to meet strength, dynamics and center-ofgravity (CG) requirements. The diagrams in figure 2 

represent a sample of the many stack sequences of 

the PM, Orbiter and Lander attempted during the 

design trade studies. The final stack configuration 

used in the design placed the Lander on top of the 

Orbiter. 

 Figure 2. Trial Launch Stack Configurations 

 The configurations placing the Orbiter on top 

were dismissed because of the large mass of the 

Orbiter and low lateral stack bending frequency that 

is created during launch. The configurations 

produced heavy structures due to the increased 

stiffness requirements necessary for raising the lateral 

bending frequency above 10 Hz. One of the design 

goals was to minimize structural mass and maintain a 

minimum frequency of 10Hz lateral and 27Hz axial. 

These values were taken from the design guide in 

reference 3. The final configuration produced the 
 

104

minimal structural mass and maintained design 

stiffness requirements. 

 

 A generic sphere-cone Lander with a mass of 400 

kg was assumed in the study and was modeled as a 

lumped mass with rigid connections to its outer 

diameter. A truss is used to create a load path from 

the Lander, through the Orbiter and into the PM. 

During the mission the Orbiter and Lander separate 

from the PM. The truss adapter to the PM is 

jettisoned with the PM and the six attachment points 

to the Orbiter heatshield are plugged. The method 

for plugging the heatshield penetrations will require 

further study. One possibility is to mechanically 

activate panels to cover the attach points. The 

Lander will separate from the Orbiter and head 

directly for the Titan surface. The upper truss 

supporting the Lander will then separate from the 

Orbiter. Aerocapture of the Orbiter will then 

commence at Titan. 

Figure 3. Stack Truss 

ORBITER AEROSHELL DESIGN 

 The Orbiter spacecraft structure consists of the 

fore body heatshield, biconic backshell, cap plate and 

internal support structure supporting the Orbiter’s 

payload. The largest payload components in the 

Orbiter are the 2.75m High Gain Antenna (HGA) and 

the spherical hydrazine tank. Figure 4 shows these 

components plus the arrangement of other internal 

components carried by the Orbiter. The payload deck 

is a hexagonal shaped aluminum honeycomb panel 

that extends to the aeroshell at six separation points. 

All payload deck items are modeled as concentrated 

masses with their CG offset made using rigid 

elements as required. The Orbiter aeroshell is 

supported by the adapter truss attached to the 

heatshield. The Lander is supported by a tube truss 

system that penetrates the Orbiter backshell at four 

locations. The load path continues straight down 

through an internal structure that also supports the 

HGA. 

 

Figure 4. Orbiter HGA and payload deck 

(Lander not shown) 

 A payload pallet ring is used to transfer the loads 

coming from the Lander as well as the payload deck 

to six hard points on the Orbiter’s heatshield. The six 

hard points are equally spaced around the perimeter 

of the payload ring and represent penetrations 

through the heatshield. The Orbiter aeroshell FEM is 

pinned at the six hard points. Concentrated masses 

were used to model the internal payloads along with 

rigid elements to properly locate CG’s. 

 The load contribution from the Lander and Orbiter 

is carried into an adapter truss through the six points 

on the heat shield. The adapter truss tapers down to 

fit the front of the PM completing the load path. The 

choice of allowing the load path to continue through 

the heat shield raises obvious concerns with the 

thermal protection system (TPS) being compromised. 

The mass of the Orbiter was reduced by not using its 

aeroshell to support the Lander mass. The stack 

concept relied on keeping load paths running axially 

through a tubular space truss. The six penetrations in 

the heatshield were accepted in this study and 

referred to as a detail requiring further investigation. 

 The Orbiter was analyzed using a combination of 

nastran finite element analysis (fea) and HypersizerTM

commercial sizing and optimization software. In 

order to utilize HypersizerTM, a coarse grid nastran 
 

105

fea was created with all non-structural masses (NSM) 

and mission loads of interest. The major NSM 

contribution was the heatshield and backshell TPS. 

Other NSM included Orbiter payload and aeroshell 

separation mechanisms as well as allocations for the 

six attach points through the heatshield. Figure 5 

shows an exploded view of the FEM used to create a 

HyperSizerTM model of the Orbiter’s aeroshell. 

Figure 5. Orbiter Aeroshell FEM 

 The mesh size was kept coarse, however included 

enough detailing of the aeroshell geometry to 

accurately calculate the element forces required in 

HypersizerTM. An opening in the backshell was 

modeled to represent an access panel to the Orbiter 

payload. No attempt was made to stiffen the opening 

by modeling the door or method of attachment to the 

backshell. The mass of the door was treated as a 

NSM with smearing at the nodes. The nastran finite 

element model (FEM) of the Orbiter was created only 

with basic nastran elements: quad4, tria3, conm2 and 

rbe2’s. These elements are easily supported by 

HypersizerTM and were imported to form a 

HypersizerTM model of the Orbiter aeroshell. 

 Several model configurations and load cases were 

used to find worse case conditions on the Orbiter 

aeroshell. The first configuration studied was the 

Orbiter in the launch mode with accelerations based 

on the Delta IV payload guide (ref. 3). The loads 

used were 3 g’s lateral and 7 g’s axial. The 

combined loads were the absolute maximums in the 

Delta IV launch load envelope. No assessment of 

acoustic energy and shock spectra on the total 

payload was attempted during the design. 

 

 The next load cases investigated were aerocapture 

entry loads of the Orbiter through the Titan 

atmosphere. Two load cases were investigated for 

different entry velocities. A 6.5 km/s and 10 km/s 

entry load cases were analyzed. The peak aero loads 

were obtained from CFD analysis based on the two 

trajectory cases. The loads were assumed to act 

normal to the heatshield and evenly distributed. The 

following loads were used: 

6.5 km/s Entry loads: 4 G axial with 3146 Pa on 

heatshield 

10 km/s Entry loads: 10.3 G axial with 8997 Pa on 

heatshield 

 The launch load forces were imported into 

HypersizerTM to start sizing of the aeroshell. 

The Orbiter’s HypersizerTM model was divided into 

different components for sizing. The approach was 

to size the heatshield as one uniform thickness as 

well as the lower and upper backshell and cap plate. 

The optimization concepts used were: honeycomb 

core with face sheets, blade stiffened panels and 

isogrids. Each concept had dimension variables that 

were used to find the optimal aeroshell geometry 

such as: blade separation distances, core and face 

sheet thickness, blade depths and thickness. 

 The TPS non-structural mass was added inside 

HypersizerTM. The mass could easily be changed on 

one of Hypersizer’sTM user input screens. This 

feature of the software was helpful for modifying the 

model to suit different TPS trial materials and 

thickness. The final TPS material used in the design 

of the Orbiter was TUFROC on the heatshield with 

an aerial density of 1.181 g/cm2. This density was 

held constant over the heatshield. The backshell and 

cap plate used SLA with an aerial density of .187 

g/cm2. The density was also constant over both 

surfaces. The TPS masses were exported as nastran 

conm2’s and evenly distributed at the element nodes. 

WHY USE HYPERSIZER? 

 Spacecraft structures contain complex geometry 

and load distributions that are highly indeterminate 
 

106

and historically demanded finite element analysis 

(FEA) to solve. Performing structural analysis and 

sizing optimization has required large degree-offreedom models with long solution run times. A 

software product called HyperSizerTM can help 

simplify structural sizing and reduce design analysis 

time. HyperSizerTM helps to automate the sizing of 

structures by reducing launch acceleration and entry 

loads into force and moment components on panels 

and beams throughout the spacecraft. The sizing 

includes finding the optimal material combinations, 

panel and beam dimensions such as thickness, depths 

and spacing. The code is not a finite element analysis 

or computer aided design package. HyperSizerTM

adds to the capabilities of these tools to allow the 

engineer to design, size and perform detailed failure 

analysis on a complete vehicle. 

 The Orbiter’s aeroshell design was used to 

demonstrate the software’s composite design 

capability and use in conceptual designs. A new 

mass-sizing tool is under development for planetary 

spacecraft at LARC. The tool will have the ability to 

link spreadsheet user inputs into HyperSizerTM for 

composite structure sizing. This will greatly improve 

structural mass estimates and lessen analysis time 

usually dominated by FEM creation and 

modifications. 

OPTIMIZATION CAPABILITIES 

 Optimization capabilities within HyperSizerTM

include finding minimum weight panel or beam 

concepts, material selections, cross sectional 

dimensions, thickness and lay-ups from a database of 

50 different stiffened and sandwich panel designs as 

well as a database of composite, metallic, honeycomb 

and foam materials. The database is used to define 

structural families inside HyperSizerTM. The 

structural families include definitions for panels and 

beams such as the “uniaxial stiffened family”, the 

“unstiffened plate/sandwich family” and the “open 

beam family”. 

 The panels shown in figure 6 below represent 

some of the typical families of structural panels 

available in HyperSizerTM. The panels may be 

stiffened with typical aerospace shapes or corrugated. 

The grid-stiffened family of panels has recently been 

added to HyperSizerTM. This allows for the sizing 

optimization of isogrids, orthogrids and general grid 

rib-stiffened panel concepts with either isotropic or 

composite materials 

 

Figure 6. Typical HyperSizerTM panels 

ORBITER PAYLOAD DECK AND HGA 

SUPPORT 

 The Orbiter payload deck and HGA support were 

modeled with plate and beam elements. The payload 

deck was a flat hexagonal shaped plate with a large 

hole cutout for a hydrazine tank carried into 

aerocapture orbit. The six corners of the hexagon 

platform extend to the outer diameter of the Orbiter. 

The platform lies in the same plane as the 

backshell/heatshield separation plane. The HGA is 

supported by an internal truss that also connects to 

the Lander truss. Loads from the Lander travel 

through the HGA support and into the payload ring 

located below the payload deck. Loads are then 

transferred through the heatshield structure and into 

the Orbiter/PM adapter. 

PROPULSION MODULE 

 The PM was modeled as shown in figure 7. The 

bulk of the module was made of aluminum channels 

and distributed lumped masses. Two solar arrays 

were also modeled and appear on the module sides. 

The solar arrays were modeled as beams having an 

approximate stiffness of the array panels. The 

modeling effort attempted to accurately capture the 

correct stiffness and mass of the module without fine 
 

107

detailing of the meshes. A concentrated mass with 

rigid elements was used for the propellant tank. 

Support structure for the tank was also provided with 

a truss system tying into a ring frame. A cylindrical 

wall stiffened with beam elements form the main 

thrust tube. The aft end of the tube was pinned with 

the forward end attaching to the Orbiter/PM adapter 

truss. The PM was modeled to help determine 

overall stack frequency at launch. By including the 

stiffness from the PM in the dynamic analysis, better 

determinations of the lowest modes were found. 

Figure 7. Orbiter FEM Components 

RESULTS 

 A dynamic analysis was performed on the Orbiter 

and launch stack to check for low natural frequencies. 

The suggested launch frequency minimums from the 

Boeing design guide (ref. 3) of 10 Hz lateral and 27 

Hz axial were used. The launch stack minimum 

modes, shown in figures 8 and 9, were 10.5 Hz lateral 

and 27.8 Hz axial respectively. Figure 10 is the first 

mode shape of the Orbiter aeroshell at 54.6 Hz. No 

dynamic magnification factors were considered 

during launch. 

Figure 8. Lateral 10.5 Hz Launch Stack Mode 

Figure 9. Axial 27.8 Hz Launch Stack Mode 
 

108

 

Figure 10. Orbiter 1st mode 

 The overall maximum deflections on the 

heatsheild were checked at launch and during entry. 

The largest deflections occurred on the heatshield 

during launch and were less than 1 mm. The 

exaggerated deflected shape of the heatshield is 

shown in figure 11. Double curvature exists where 

the six adapter truss points attach through the 

heatshield. The deflections were small and not 

considered a concern for TPS bonding. 

Figure 11. Orbiter Deflections at Launch 

 The Orbiter’s aeroshell was sized after several 

iterations between nastran and HypersizerTM. The 

dominant load case was found to be during launch. 

The process optimized the aeroshell structure and 

indicated which materials and structural concept 

would produce the lightest aeroshell. Honeycomb 

core with face sheets were shown to be the best 

structural concept. The final core material for the 

heatshield was a 25.4mm thick Hexcell 5052 alloy 

hexagonal aluminum honeycomb with 1.7mm 

graphite polyimide face sheets. This design was 

similar to the Mars Exploration Rover (MER) 

heatshield except for the six hard points used in 

attaching the Orbiter heatshield to the PM. The 

backshell sizing done within HyperSizerTM showed a 

honeycomb core face sheet concept produced the 

minimum structural mass. The core was a 12.7mm 

thick Hexcell 5052 alloy hexagonal aluminum 

honeycomb and graphite polyimide face sheets of 

varying thickness. The cap plate design was similar 

to the backshell. 

 Results from analyzing the HGA and Lander 

trusses showed the optimal material was 2” OD, 

0.12” wall M55J/954 tubes. The Orbiter/PM adapter 

truss was similarly made with 3.2” OD, 0.2” wall 

M55J/954 tubes. The sizes were driven by finding 

sections large enough to prevent buckling. 

 A summary of the final Orbiter aeroshell mass is 

given in table 1 and the total launch stack mass 

summary is shown in table 2. The total spacecraft 

mass for the launch configuration was 3173.2 kg and 

included the Lander, Orbiter and PM. TPS and nonstructural masses were included plus allowances for 

miscellaneous items such as heatshield to backshell 

separation components. The design relies on a 

system of composite M55J trusses that form a load 

path into the PM. This system produced minimal 

displacements during launch and held stresses within 

safety limits that were: 1.4 on ultimate, 1.25 on yield 

limit and 1.5 for buckling. 

Part Area Structure 

Mass 

TPS 

Mass 

NSM 

Heatshield 

(TufRoc) 

12.58 m2 41.58 kg 148.62 kg 0 

Backshell 

(SLA-561V) 

15.01m2 43.27 kg 28.69 kg 2.38 kg 

Pallet Ring 1.20m2 42.47 kg 0 1.20 kg 

Separation 

Ring 

1.79m2 11.35 kg 0 .89 kg 

Separation 

Ring 

Attachments 

.45m2 2.85 kg 0 4.50 kg 

Totals 141.52 kg 177.31kg 8.97 kg

Total Aeroshell (structure + TPS + NSM) = 327.80 kg

Table 1. Orbiter Aeroshell Mass 

Lander 400 kg 

Lander Truss 61.8 kg 

Orbiter Aeroshell + 

Payload 

1200 kg 

Orbiter/PM Truss 61.4 kg 

PM 1450 kg 

Total spacecraft 3173.2 kg 

Table 2. Launch Stack Mass 

CONCLUSION 
 

109

 The success of an aerocapture mission at Titan 

greatly depends on the mass reduction of the 

structure and the configuration of the launch stack. 

The design efforts encountered during this conceptual 

study showed the importance of defining the 

configuration in reducing spacecraft mass. The final 

launch configuration used an unconventional method 

of attaching the Orbiter heatshield to the PM. This 

method allowed a continuous load path from a 400kg 

Lander, into the Orbiter, through the Orbiter 

heatshield and into the PM. Maintaining a load path 

through the trusses that avoided the Orbiter aeroshell 

from supporting the Lander minimized the Orbiter 

aeroshell. The stacking arrangement also minimized 

the buckling lengths of the truss members as well as 

the number of required members. HyperSizerTM was 

used to perform optimization sizing of the Orbiter 

aeroshell without a detailed mesh and extensive 

remodeling effort. The results indicated a 

honeycomb face sheet composite could produce a 

light structure while providing the necessary stiffness 

to meet minimum dynamic frequency requirements at 

launch. The results from this study have established 

a starting point for a detailed fea of the Orbiter 

aeroshell. Such an analysis could include varying 

core and lay-up thickness and detailed analysis of 

attachment connections and separation mechanisms. 

The structural mass for this design was within the 

mass margin estimated for a successful Titan 

mission. 

ACKNOWLEDGEMENTS 

 The author recognizes the following for their 

contribution to the Titan structural analysis: 

Jonathan Lam (JPL) for his nastran modeling of the 

PM and Orbiter payload structure, Rob Bailey (JPL) 

for packaging the Orbiter and detailing payload 

requirements, Bernie Laub (ARC) for TPS sizing on 

the Orbiter heatshield and backshell, and 

 Eric Dyke (Swales Aerospace-LARC) for 

investigation of the Orbiter/Lander truss and stacking 

configurations. 

REFERENCES 

1. G. A. Hrinda, “Integrated Structural Analysis of 

the X-43 Using HyperSizer Sizing Software”,

CS/PSHS/APS Joint Meeting, November 2000. 

2. G. A. Hrinda, “Structures For the 3rd Generation 

Reusable concept Vehicle”, Space 2001 Conference 

 

3. “Delta IV Payload Planners Guide”, The Boeing 

Company, Huntington Beach, CA, 2000. 

4. Collier Research Corporation, “HyperSizer 

Tutorial & Applications”, 2nd edition, Collier 

Research Corporation, 1998. 

5. C. S. Collier, P. W. Yarrington and M. 

Pickenheim, “Design Optimization Using 

HyperSizer”, Collier Research Corporation, The 1998 

MSC Americas Users 'Conference’, Universal City, 

CA, Oct 5-8, 1998. 

6. C. S. Collier, P. W. Yarrington and M. 

Pickenheim, “Next Generation Optimization Today”, 

The 1997 MSC Aerospace Users Conference', 

Newport Beach, CA, November 18-20, 1997.
110

PLANETARY PROBE MASS ESTIMATION TOOL DEVELOPMENT AND ITS 

APPLICATION TO TITAN AEROCAPTURE 

R. Eric Dyke 

Swales Aerospace 

Hampton, VA 23681 

ABSTRACT 

 An integrated mass properties estimation and 

aeroshell structural sizing design tool, called 

PROBECODES, is being developed for quick sizing of 

conceptual, atmospheric-entry planetary exploration 

vehicles. This tool will eventually integrate EDS PLM 

Solutions I-deas® entry system geometry models, 

Microsoft® Excel-based entry system mass properties 

estimation spreadsheets, CFD generated 

aerothermodynamic loads, and structural 

sizing/optimizing software such as EDS PLM Solutions 

I-deas® or Collier Research’s HyperSizerTM in order to 

create preliminary vehicle mass estimates, with 

particular emphasis on aeroshell structural sizing 

estimates. The Microsoft® Excel-based spreadsheet 

portion of the tool, called ProbeMAASS1 is currently 

being developed. It contains a mass properties/design 

database of previous planetary exploration vehicles and 

design studies; a mass properties breakdown and design 

description of these vehicles’ major subsystems; mass 

estimation methods for the major subsystems of new 

vehicles using combinations of analytical techniques, 

empirical relationships, and user-specified data; and an 

automatic aeroshell structural finite element mesh 

generator. Currently this tool is limited to sphere-cone 

and ellipsled type vehicles, but future versions will 

include other types as they are developed. Though still 

under development, parts of this tool were successfully 

used to help quickly estimate Orbiter aerocapture 

system masses for a wide variety of TPS 

material/thickness and aeroshell size combinations for 

the Titan Aerocapture Systems Analysis study of 2002. 

In addition this tool was used to help estimate total 

launch wet mass for ten different Titan mission/launch 

vehicle configurations. This paper will present an 

overview of the full tool (PROBECODES) and an 

overview of the mass properties estimation spreadsheets 

(ProbeMAASS1). Lastly, it will present the mass 

sensitivity results from the Titan Aerocapture study. 

NOMENCLATURE/ABBREVIATIONS 

CBE Current Best Estimate 

CFD Computational Fluid Dynamics 

CG’s Centers of Gravity 

EGA Earth Gravity Assist 

GDTL Geometric Data Transfer Link 

GUI Graphical User Interface 

L/D Lift divided by Drag 

LMDTL Lumped Mass Data Transfer Link 

LV Launch Vehicle 

MEL Master Equipment List 

MEM’s Mass Estimation Methods 

MPD Mass Properties Database 

MPB Mass Properties Breakdown 

SEP Solar Electric Propulsion 

TPS Thermal Protection System 

VVVGA Venus-Venus-Venus Gravity Assist 

VGA Venus Gravity Assist 

∆V Velocity change 

INTRODUCTION 

 As part of the ongoing exploration of the solar 

system, there is a need for an atmospheric-entry probe 

design tool that takes conceptual probe design data and 

provides quick-turnaround probe mass estimates and 

aeroshell structural sizing. To fulfill this need, an 

integrated mass properties estimation and aeroshell 

structural sizing design tool is currently being

developed. This tool, called PROBECODES (Probe 

Conceptual Design Software), is comprised of several 

different components, including commercially available 

software, as shown in Figure 1. Current methods for 

probe mass estimation and aeroshell sizing often 

involve using separate, non-integrated historical data, 

mass estimation methods, and finite element 

modeling/structural sizing tools. Often the geometry 

models, mass estimations, and finite element modeling 

are done by different individuals using different tools 

which may or may not share data easily. The ultimate 

goal of PROBECODES is to integrate mission-driven 
111

payload requirements, EDS PLM Solutions I-deas®1 

entry system geometry models, Microsoft® Excel2-

based entry system historical databases, Microsoft®

Excel-based mass properties estimation spreadsheets, 

CFD-generated aerothermodynamic loads (pressure 

and/or temperature distributions), structural materials 

data, and structural sizing/optimizing software such as 

MSC/NASTRAN3

, EDS PLM Solutions I-deas® or 

Collier Research’s HyperSizerTM 4 into a seamless endto-end tool in order to produce preliminary vehicle 

mass and aeroshell structural sizing estimates. The 

primary benefit of PROBECODES, when completed, 

over previous methods will be to allow a single user to 

quickly create conceptual probe geometry, estimate 

component masses, and size the aeroshell structure all 

within a single tool in a short time. 

 PROBECODES and the mass estimation 

component, ProbeMAASS1 (Probe Mass 

Approximations & Automeshing for Structural Sizing, 

version 1), are currently limited to sphere-cone and 

ellipsled shaped entry vehicles. Future entry vehicle 

shapes, such as bi-conics, bent bi-conics, and other 

configurations will be addressed in later versions. This 

tool also assumes a general payload packaging layout 

has been established. Lastly, the automated structural 

sizing aspects of the tool are currently limited to the 

aeroshell. Sizing for the wide variety of possible 

internal structural components such as payload support 

structure (decks, rings, etc.) is not currently supported. 

Such items must be handled in an “offline” manner, 

with the resulting finite element meshes and masses 

then being used as input to this tool. 

 While still in the development stages, portions of 

ProbeMAASS1 were successfully used for mass 

sensitivity studies for the Titan Aerocapture Systems 

Study of 20025

. Aeroshell mass sensitivities were 

developed for a range of provided TPS materials and 

aeroshell size/TPS material combinations. In addition, 

the tool was used to help estimate total (full stackup) 

launch wet masses for ten different mission 

architectures. The results of theses studies are 

presented later in this paper. 

OVERVIEW OF PROBECODES TOOL 

 The baseline requirements in developing 

PROBECODES are that preliminary aeroshell 

structural design and probe mass estimate results be 

produced from the conceptual data in a relatively short 

time (2 to 3 days), and that these results be produced 

with a minimum of special hardware and software. The 

current tool architecture primarily uses EDS PLM 

Solutions I-deas®, Microsoft® Excel, and Collier 

Research’s HyperSizerTM. I-deas® is used for aeroshell 

solid modeling and surface area/volume calculations 

and for aeroshell structural finite element model editing 

and solving. Microsoft® Excel is used for the probe 

mass properties database of historical probes and probe 

studies, the mass estimation methods, aeroshell surface 

area calculations, and aeroshell structural finite element 

model generation. Though not yet integrated into 

PROBECODES, Collier Research’s HyperSizerTM will 

be used to optimize the aeroshell structure. As newer, 

better methods become available, or as the tool 

components allow themselves to be combined or 

separated, the tool architecture as shown in Figure 1 

will be modified to incorporate those changes to 

produce a more efficient product. 

 

PAYLOAD REQUIREMENTS, MISSION DESIGN 

 The user must start with a given or assumed mission 

profile. The mission profile determines the payload 

components required to meet the mission objectives, 

including scientific instrumentation and their associated 

support subsystems (telecom, navigation, command and 

data handling, power, etc.). The mission profile also 

determines the delivery time and atmospheric entry 

profile, including entry speed, entry angle, maximum 

deceleration, and aeroheating rates and total heat loads. 

The aerodynamic performance requirements, such as 

vehicle L/D, ballistic coefficient, etc., necessary to 

ensure the proper entry profile help determine the 

general class of entry vehicle. Currently, the low L/D 

vehicles are represented by the axisymmetric spherecone shapes, and the mid L/D vehicles are represented 

by the ellipsled shape. Together, the general vehicle 

shape required and the volume necessary to 

accommodate the proposed payload determine the 

overall aeroshell size. 

SOLID MODELING (EDS PLM SOLUTIONS IDEAS®) 

 EDS PLM Solutions I-deas® is used as the entry 

vehicle solid modeler. Multiple axisymmetric spherecone solid models have been created within an I-deas®

model file, including a stand-alone forebody, standalone backshells (conic, bi-conic, multi-conic, 

hemispheric, and spherical cap), and combined 

forebody/backshell models for bi-conic and 

hemispheric cap backshells. These models are 

parametrically dimensioned such that the user may edit 

one or all of the cross section dimensions and the solid 

model is automatically updated. An axisymmetric 

ellipsled solid model has also been created with a single 

backshell separation configuration. Additional solid 

models will be created for more general (nonaxisymmetric) ellipsled shapes with more general 

backshell separation geometries. 
112

GEOMETRIC DATA TRANSFER LINK (GDTL) 

 The GDTL transfers the solid model geometry data 

from I-deas® to the ProbeMAASS1 spreadsheets. With 

the appropriate solid model opened, an I-deas® program 

file is used to list and save the model dimensions and 

surface areas. The program file is then transferred (if 

necessary) to the PC platform where ProbeMAASS1 

resides. A macro within ProbeMAASS1 then reads the 

file and uploads the dimensions and surface areas for 

use in later calculations. 

ENTRY SYSTEM DATABASE 

 The entry system database is a compilation of 

planetary exploration probe and probe study design 

data. It is in Excel spreadsheet form, with separate 

sheets for the sphere-cone and ellipsled vehicle types. 

These databases, described in more detail in a later 

section, are part of the ProbeMAASS1 spreadsheets, 

and contain probe geometry, entry environment 

definition, subsystem and component masses, and 

structural materials and construction methods. The data 

are used to develop mass estimation methods in 

ProbeMAASS1, and allow the user to make quick 

initial guesses for subsystem or component masses 

based on similarity to previous designs. 

USER INPUT PARAMETERS 

 User input parameters are values entered into the 

ProbeMAASS1 spreadsheets. They include such wideranging details as probe destination, number of 

propellant tanks, or parachute diameter. They are input 

in a “linear” fashion on each spreadsheet from top to 

bottom, as each spreadsheet is filled out for each probe 

system. ProbeMAASS1 uses the historical database, 

user input, and mass estimation methods (MEM’s) 

discussed later to calculate component masses. The 

user can also override any or all calculated values in the 

spreadsheets. Since the ProbeMAASS1 spreadsheets 

will calculate aeroshell surface areas that are used in 

aeroshell structure and TPS mass estimation, the user 

may also input probe dimensions separate from the 

imported I-deas® geometry. This allows the user to 

take quick looks at the impact dimension changes have 

on aeroshell mass. 

ENTRY SYSTEMS MASS PROPERTIES TOOL 

(PROBEMAASS1) 

 ProbeMAASS1 is the heart of this mass estimation 

tool. It will be discussed in more detail in a later 

section. 

MASS PROPERTIES OUTPUT 

 The mass properties output currently consists of 

three items: the lumped mass data, the aeroshell mesh, 

and the probe mass summary. The lumped mass data 

consists of point masses and their associated centers of 

gravity (CG’s) for larger components. They are either 

calculated by the ProbeMAASS1 spreadsheets or 

entered directly by the user, and are added to the 

aeroshell mesh file. While the spreadsheets currently 

determine the point masses, the CG calculations have 

not yet been fully integrated into ProbeMAASS1. The 

aeroshell mesh is a structural finite element mesh 

created by internal macros in the ProbeMAASS1 

spreadsheets and is described in more detail in a later 

section. The probe mass summary sheet lists current 

best estimate (CBE) masses for each probe component. 

Component masses are also added to give associated 

subsystem masses and total probe mass. The user may 

also specify separate uncertainty factors to be applied to 

each component CBE to determine the growth mass. In 

the first pass through ProbeMAASS1, before aeroshell 

structural sizing has been completed, the aeroshell mass 

is calculated parametrically from historical probe data. 

Due to the wide variety of probe configurations, this 

first aeroshell mass estimate is considered low fidelity 

and must be updated after the structural sizing has been 

completed on the aeroshell. 

INERTIAL, AERO & THERMAL LOADS 

 These loads are determined by external analyses. 

Simple inertial loads (gravity and body accelerations) 

and uniform aerodynamic pressures can be applied to 

the mesh within the ProbeMAASS1 spreadsheets and 

macros. Thermal loads/temperature distributions and 

more complex aerodynamic pressure distributions must 

be applied within I-deas® or some other pre/post 

processor at this time. 

MATERIAL DATA 

 The material data for finite element analysis is 

typically specified by the user within a finite element 

pre/post processor. Since the aeroshell structure is to be 

optimized using Collier Research’s HyperSizerTM (see 

below), the current tool architecture assumes the 

aeroshell will be optimized using a HyperSizerTM

material database. 

FINITE ELEMENT MODELING/STRUCTURAL 

SIZING 

 The aeroshell mesh is imported into I-deas® or other 

appropriate pre-processor. There the mesh is 

edited/modified, lumped masses are added, and 
113

secondary or payload support structure is added as 

necessary. The model is then solved in I-deas® or 

MSC/NASTRAN. The mesh input and results files are 

then submitted to HyperSizerTM for structural 

optimization of the aeroshell. 

 HyperSizerTM is a windows-based, commercially 

available structural sizing/weight estimation/weight 

optimization software tool4

. A plate and/or beam 

element finite element model and its results file are first 

read into HyperSizerTM. A graphical interface allows 

the user specify a range of structural “families” (i.e., 

unstiffened panels, sandwich panels, uniaxially 

stiffened panels), a range of structural “concepts” (i.e. 

“C” or “I” stiffeners), and a range of materials for 

various regions of the finite element model. 

HyperSizerTM then steps through each combination of 

designs for each region from lightest to heaviest until a 

successful design is created. The major advantage of 

HyperSizerTM is that it performs numerous user 

specified closed-form failure checks on each region for 

each design (family/concept combination) using the 

initial finite element model. This eliminates the need to 

create multiple finite element models. 

 

 Once aeroshell optimization is complete, the 

aeroshell mass is entered into the probe summary 

spreadsheet. Future enhancements will include 

automatic updates of the probe mass summary 

spreadsheet after aeroshell optimization. 

STRUCTURAL DESIGN AND MASS ESTIMATE 

(PRELIMINARY) 

 As part of the HyperSizerTM optimization routine, 

several structural concepts can be evaluated, including 

sandwich construction, blade stiffened panels, isogrids, 

etc. Thus, an aeroshell mass and structural design are 

produced as part of the HyperSizerTM optimization 

process. 

OVERVIEW OF PROBEMAASS1 

 ProbeMAASS1 is the main component of 

PROBECODES. It is a series of Excel spreadsheets 

which act as the “clearing house” for all probe design 

data, and is the primary user interface for probe mass 

estimation. It contains an import interface for aeroshell 

solid model dimensions and surface areas created in Ideas®. It also contains Visual Basic macros for 

generating aeroshell structural finite element meshes for 

sphere-cone and ellipsled geometries. The approach in 

developing ProbeMAASS1 was to use a linear 

progression through the series of Excel spreadsheets 

with a minimum of user inputs in order estimate the 

entry probe mass. Early on in the tool development, the 

Excel spreadsheet format was chosen over various 

“GUI” formats so that all historical data, references, 

supporting notes, and equations could easily be viewed 

by the user. Figure 2 shows the major functional 

features of ProbeMAASS1. These features are 

discussed in more detail in the following sections. 

MASS PROPERTIES DATABASE (MPD) 

 The Mass Properties Database (MPD) is a 

collection of historical probe (flight) and probe study 

(paper) data. It contains a limited amount of data for 

European, Soviet, and Japanese probes, but is primarily 

a collection of data from NASA-directed probes. There 

are separate database spreadsheets for the sphere-cone 

and ellipsled type entry vehicles. The probe data are 

grouped according to probe destination (i.e. Venus, 

Mars, etc.), and are subdivided into the major 

categories shown in Figure 3. The first four categories 

are probe overviews: Mission Overview (destination, 

launch date, launch vehicle), Total Mass Properties 

(mass, moments of inertia), Geometry (overall forebody 

and backshell dimensions), and Entry Environment 

(speed, angle, maximum deceleration, maximum heat 

rate and heat load). Next come the probes major 

systems: Forebody System, Aftbody System, 

Deceleration Systems, and Payload. Each system is 

further broken down into the subsystems shown in 

Figure 3, which are described in more detail in the next 

section. This general breakdown is followed for both 

sphere-cone and ellipsled type entry vehicles. There is 

considerably less data in the ellipsled database simply 

due to the limited amount of ellipsled flight and study 

data available. 

 As with all other spreadsheets in ProbeMAASS1, 

the database spreadsheets follow a color-coding 

convention. For the database sheets, cells containing 

flight data are colored turquoise, while cells containing 

paper study data are colored light turquoise. Some 

values, like probe ballistic coefficient, which are not 

found specifically in the literature, are calculated from 

other data, and the resulting cells are colored tan. Once 

the color conventions are understood, they help the user 

understand at a glance the type of data he or she is 

dealing with. 

 Every cell or every line of cells in the MPD contains 

a cell comment that lists a reference for the data shown. 

For cases where no data was found, the associated cell 

was left blank. In other cases, conflicting values were 

found, such as for a heat load or entry deceleration. In 

such cases, the most recent data value was entered into 

the cell, and other conflicting values, with references, 

were listed in the cell comment. Cell comments are 

also used to include more in-depth information beyond 
114

what is showable in the given cell, such as facesheet or 

honeycomb thicknesses for sandwich panel aeroshells. 

MASS PROPERTIES BREAKDOWN (MPB) 

 The probes’ major system masses are further refined 

by the Mass Properties Breakdown (MPB). The MPB 

is not a separate spreadsheet, but is embedded in and 

part of the MPD. The actual refinement is also shown 

in Figure 3. For the forebody and aftbody systems, the 

probe mass is broken down to thermal protection, 

structure, separation system, thermal control, and 

miscellaneous subsystems and their components. The 

deceleration systems are further broken down into the 

pilot, main, and terminal parachutes and terminal 

descent (airbag) subsystems. Even within these 

subsystems is more detailed information such as 

parachute deploy speed, deploy altitude, etc. The 

payload system is further broken down into structure, 

science instrument, power, guidance/navigation, 

telecom, command and data handling, thermal control, 

harness, propulsion, and miscellaneous subsystems. 

Again, within these subsystems is considerable detailed 

component mass such as propellant tank size and 

material, etc. Some of the more detailed information is 

shown in the cells, while other is contained in the cell 

comments. 

MASS ESTIMATION METHODS (MEM’S) 

 The bulk of ProbeMAASS1 is devoted to the Mass 

Estimating Methods (MEM’s). These are the individual 

spreadsheets, as shown schematically in Figure 4, 

which are used to estimate the component, subsystem, 

system, and total probe mass. A separate spreadsheet is 

devoted to each major system and its associated 

subsystems and components. The MEM’s take several 

forms, depending on the system or component being 

estimated. In some cases, the user may simply refer to 

the MPD and use an historical component mass which 

is most similar to the current design, and which may be 

independent of probe size. An example would be 

aeroshell separation fittings. While the number of 

fittings may vary, the basic fitting design may be 

relatively constant, thus allowing the user to apply the 

historical mass to the current design for a quick mass 

estimate. 

 In other cases, simple parametric scaling is used to 

estimate a new component mass. For example, a new 

parachute mass may be estimated by scaling a closely 

similar historical parachute mass by the ratio of the 

diameter squared. Such scaling assumes a constant 

areal density (kg/m2

) between the historical design and 

the new design. Areal densities are also used to 

estimate masses for aeroshell structure and TPS. Using 

the aeroshell surface areas either from the geometry 

imported from the I-deas® solid model or calculated by 

the ProbeMAASS1 spreadsheets, in conjunction with 

areal densities from historical data, the aeroshell 

structure and TPS masses can be calculated. The 

historical areal densities are generally curve fits of 

previous probe design data. Figure 5 shows an example 

of a curve fit used to give a first-pass estimate of 

forebody aeroshell structural areal density for spherecone type probes. 

 The highest fidelity mass estimation uses closed 

form solutions. For this tool, closed form solutions are 

primarily used to calculate sphere-cone surface areas, 

and to estimate mass for the propulsion system and its 

components. Standard textbook equations 6

 are used to 

estimate propellant mass, pressurant mass, and 

propellant and pressurant tank masses. While this 

closed form method may produce higher fidelity mass 

estimates, it requires more user input than the other two 

methods. 

 As with any tool, care must be used to ensure that 

correct and reasonable input is being used, and that the 

output is reasonable. While historical masses may be 

used for some components, technological advancements 

often reduce the size and mass of these components. 

Such mass reductions are not currently predicted in the 

MEM’s. In addition, several of the MEM’s in this tool 

use curve fitting from only a few historical data points 

to develop scaling parameters. Such curves are often 

based on a very limited number of data points. The 

MEM’s are updated to incorporate new data as it 

becomes available. To that end, this tool has been 

designed with a great deal of flexibility to allow the 

user to override estimated masses at any level in the 

process (component, subsystem, full probe). 

 Validation of the MEM’s is an ongoing process. 

Closed form solutions for sphere-cone surface areas and 

propulsion component masses have been checked by 

hand to verify their accuracy. The surface area 

calculations for ellipsled geometry using the internal 

macros have been compared with I-deas® calculated 

values and match to within 0.05%. Curve fits of 

historical data and parametric mass scaling laws are 

constantly being reviewed to ensure the most accurate 

mass estimation methods. 

AEROSHELL AUTOMESHER 

 The last major feature of ProbeMAASS1 is the 

automatic aeroshell structural finite element mesh 

generator and its associated surface area calculators for 

the sphere-cone and ellipsled aeroshells. The spherecone probe dimensions and surface areas can be 
115

imported through the GDTL discussed earlier. In order 

to provide more flexibility and to quickly evaluate the 

impact dimensional changes have on probe mass, the 

user may also enter override dimensions for the 

forebody and aftbody in their respective spreadsheets. 

Figure 6 shows generic sphere-cone forebody and 

aftbody cross-sections and dimensions. The 

spreadsheets then calculate surface areas using standard 

equations 7. The I-deas® or override dimensions are 

then carried into the Sphere-Cone Aeroshell Mesh 

Generator spreadsheet (See Figure 4.) where the user 

then enters values to control the mesh density. Once 

all mesh control variables are entered, the internal 

macro generates a plate element structural finite 

element mesh with dummy material properties. The 

user may specify whether the forebody only, backshell 

only, or combined forebody and backshell meshes are 

created. For the sphere-cones, there is only one 

separation configuration: the backshell separating from 

the forebody at a given horizontal plane. The spherecone macro supports forebodies with hemispherical, 

spherical cap, conic, bi-conic, or multi-conic 

backshells. Figure 7 shows a typical mesh for a spherecone aeroshell with forebody and bi-conic backshell. 

The resulting mesh file is output as an 

MSC/NASTRAN .dat bulk data file. This format 

allows for easy model import into a wide variety of 

finite element pre/post processors (FEMAP, PATRAN, 

I-deas®, etc.). After import into the pre-processor, the 

user may edit the model as necessary to incorporate 

lumped masses and/or secondary structure or modify 

material properties, then solve the model. The user 

then sends the model and solution files to HyperSizerTM

for structural optimization. Currently, the link to 

HyperSizerTM is a manual transfer. Future 

enhancements will include automated links. 

 For ellipsled entry vehicles, dimensions and surface 

areas can also be brought in through the GDTL. At this 

time, the GDTL only supports axisymmetric ellipsled 

vehicles with horizontal aft and canted forward 

backshell/forebody separation configurations. The 

GDTL will be updated later to transfer geometry data 

for more general ellipsled configurations. The 

ProbeMAASS1 Ellipsled Surface Area Calculator 

spreadsheet currently allows the user to enter 

dimensions for more general ellipsled geometries. 

Figure 8 shows a generic ellipsled with associated 

dimensions. Ellipsleds may be non-axisymmetric 

(upper and lower halves with different ellipse semi-axis 

ratios), and may have one of eighteen different 

backshell separation configurations. Similar to the 

sphere-cones, the dimensions here are carried into the 

Ellipsled Aeroshell Mesh Generator spreadsheet where 

the user enters all the required mesh control values, and 

the internal macro creates the aeroshell plate element 

structural mesh. The mesh control values allow the 

user to specify creation of the ellipsled forebody only, 

backshell only, base only, or all three. Figure 9 shows a 

typical ellipsled plate element mesh generated by the 

automesher. The resulting mesh can then be imported 

into a pre/post processor for editing and solving, then 

sent to HyperSizerTM for structural optimization. 

FUTURE ENHANCEMENTS 

 ProbeMAASS1 is continually being updated. As 

new historical data becomes available, it is added to the 

database. This new data is then used to update the 

MEM’s where applicable. Currently, the automatic 

mesh generators only create plate elements for the 

aeroshell. Future enhancements include a lumped mass 

data transfer link (LMDTL). Within the spreadsheets, 

centers of gravity (CG’s) will be estimated for various 

point masses such as fuel tanks, electronics boxes, 

parachutes, etc. The user will also be able to override 

these estimates by inputting the desired CG’s. The 

point masses and their associated CG’s will be added to 

the finite element model when it is generated. In 

addition, more automated links are planned to open and 

transfer geometric data from the I-deas® solid model 

files into the ProbeMAASS1 spreadsheets, and to 

access and run the structural finite element codes such 

as I-deas® and HyperSizerTM. 

APPLICATION TO TITAN AEROCAPTURE 

SYSTEMS STUDY (2002) 

 Though still under development, parts of this tool 

were successfully used in conjunction with other 

subsystem mass estimates8 provided by the Jet 

Propulsion Laboratory (JPL) and TPS sizing9,10,11,12

provided by Ames Research Center (ARC), Langley 

Research Center (LaRC), and Applied Research 

Associates (ARA) to help estimate Titan orbiter masses 

for a wide variety of TPS material/thickness and 

aeroshell size/TPS combinations and to help quickly 

estimate total wet launch masses for the Titan 

Aerocapture Systems Analysis study of 2002. 

 IMPACT OF TPS SELECTION 

 The goal of the TPS selection sensitivity study was 

to determine the impact of aeroheating levels and, 

hence, TPS material selection and thickness on the total 

Titan orbiter mass, the aerocapture system mass, and 

the orbiter forebody TPS mass. In addition, the impact 

on aerocapture system mass fraction and orbiter 

ballistic coefficient were also investigated. This study 

required separate detailed aeroheating and TPS analyses 

for a 6.5 km/s Titan atmosphere entry to determine 

aeroheating rates and loads (both radiative and 
116

convective), and TPS material selection and thickness. 

For this study, several assumptions were made: 

1. The orbiter aeroshell was held at a 

constant 3.75 meter maximum diameter. 

2. The backshell TPS material (SLA) and 

areal density (1.87 kg/m2

) were held 

constant9

. 

3. The orbiter wet mass (without aerocapture 

system) 8

 was held constant. 

4. The aeroshell structural areal density 

(kg/m2

) 13 was held constant. 

5. The forebody TPS thicknesses and areal 

densities, shown in Table 1 below, were 

held constant for each case. All values 

except C-C data are from References 9-

11. C-C data is from Reference 12. 

Material Thermal Load 

Case 

Thickness 

(cm) 

Areal 

Density 

(g/cm2

) 

SRAM-14 Nominal lift down 1.55 0.348 

SRAM-17 Nominal lift down 1.93 0.526 

SLA-561V Nominal lift down 2.43 0.622 

SRAM-20 Nominal lift down 2.08 0.667 

SRAM-20 0.8 conv, 1.8 rad 2.54 0.814 

SRAM-20 0.6 conv, 2.6 rad 3.76 1.204 

PhenCarb-20 Nominal lift down 2.34 0.711 

PhenCarb-20 0.8 conv, 1.8 rad 2.71 0.868 

PhenCarb-20 0.6 conv, 2.6 rad 2.91 0.931 

TUFROC Nominal lift down 5.13 1.181 

TUFROC 0.8 conv, 1.8 rad 5.70 1.289 

TIFROC 0.6 conv, 2.6 rad 6.12 1.371 

C-C (Genesisstyle) +15% Nominal lift down 5.51 1.493 

C-C (Genesisstyle) +15% 0.8 conv, 1.8 rad 5.99 1.587 

C-C (Genesisstyle) +15% 0.6 conv, 2.6 rad 6.39 1.665 

Table 1. Titan Aerocapture TPS Sizing Data 

 The nominal orbiter dimensions were adjusted for 

each material and thickness in order to calculate surface 

areas at the mid-thickness of the forebody TPS. These 

dimensions were entered into the Sphere-Cone 

Forebody spreadsheet in ProbeMAASS1 (See Figure 

4.) as user input, and the spreadsheet calculated the 

associated surface areas. TPS areal densities were also 

entered later as user input, and the TPS mass was 

calculated by the spreadsheet. Fifteen TPS 

material/thickness combinations were evaluated and 

TPS masses calculated in just a couple hours. The 

resulting masses were added to other previously 

calculated orbiter system masses to determine total 

orbiter wet launch mass, aerocapture system mass, 

aerocapture system mass fraction, forebody TPS mass, 

and aerocapture ballistic coefficient as a function of 

TPS. The above items were also calculated for TPS 

with 30% and 50% mass margin. The results were 

plotted as bar graphs14 as shown in Figures 10 through 

14. The TPS material TUFROC was chosen as the 

baseline. ProbeMAASS1 allowed the user to show the 

orbiter mass sensitivity for a wide variety of TPS 

materials and thicknesses in a very short time. For this 

particular study, the lightest candidate investigated 

(SRAM-14) showed a potential 128 kg mass savings 

over the baseline TUFROC TPS material. 
117

IMPACT OF ORBITER DIAMETER/TPS 

COMBINATION 

 The goal of the orbiter diameter/TPS combination 

sensitivity study was to determine the impact of 

aeroshell diameter/forebody TPS combination on total 

Titan orbiter mass, aerocapture system mass and mass 

fraction, orbiter forebody TPS mass, and orbiter 

aerocapture ballistic coefficient. Several assumptions 

were made for this study: 

1. The 3.75 meter diameter aeroshell was 

photographically scaled down by 5%, 

10%, 15%, and 20%. Payload packaging 

was not re-evaluated. 

2. The backshell TPS material (SLA) and 

areal density (1.87 kg/m2

)

9

 were held 

constant. 

3. All Titan orbit dry mass subsystem 

masses were held constant except 

structure, which was reduced linearly with 

diameter. 

4. The aeroshell structural areal density 

(kg/m2

)

13 was held constant. 

5. The orbiter separation ring and payload 

pallet ring structure masses13 were scaled 

linearly with diameter. 

 Similar to the study of mass sensitivity to TPS 

material alone, the modified aeroshell dimensions and 

TPS areal densities were input into the ProbeMAASS1 

spreadsheet which calculated TPS masses. The TPS 

masses and orbiter systems masses scaled per items 4 

and 5 above were added to other previously calculated 

orbiter system masses to determine total orbiter wet 

launch mass, aerocapture system mass, aerocapture 

system mass fraction, forebody TPS mass, and 

aerocapture ballistic coefficient for the various 

diameter/TPS combinations. The results for the five 

aeroshell diameters and six TPS materials under 

nominal aeroheating loads were plotted as bar graphs14

as shown in Figures 15 through 19. The 3.75 meter 

diameter aeroshell with TUFROC TPS material was 

chosen as the baseline. Again, ProbeMAASS1 allowed 

the user to show the orbiter mass sensitivity for a wide 

variety of aeroshell diameter/TPS combinations in a 

very short time. This particular study showed the 

potential mass savings associated with reducing 

aeroshell diameter and showed significant mass 

variations with respect to aeroshell diameter and TPS, 

with the 3.0 m aeroshell/SRAM-14 TPS combination 

having 19.2% of the TPS mass of the baseline 3.75 

m/TUFROC combination. 

IMPACT OF MISSION 

CONFIGURATION/LAUNCH VEHICLE ON 

TOTAL LAUNCH GROWTH WET MASS 

 The goal of this sensitivity study was to determine 

the impact mission configuration/launch vehicle had on 

total launch growth wet mass. Several assumptions 

were made for this study: 

1. Two all-chemical missions and eight 

aerocapture mission architectures, with 

either chemical or SEP (Solar Electric 

Propulsion) cruise propulsion systems15, 

were evaluated. See Table 2 below. 

2. For the aerocapture missions, a constant 

515 kg aerocapture system growth mass 

and a constant 875 kg payload growth 

mass (in Titan orbit) was assumed. 

3. For the all-chemical missions, a constant 

937 kg payload growth mass (in Titan 

orbit) was assumed. 

4. The EGA mission with SEP and 

aerocapture and a Delta 4450 launchvehicle was the baseline. 
118

Table 2. Mission Profiles15

 ProbeMAASS1 was used to a lesser extent for this 

sensitivity study, but still proved valuable for producing 

the desired mass estimations in a very short time. For 

this study, the payload mass was assumed a constant 

1390 kg for the aerocapture missions, and a constant 

937 kg for the all-chemical missions. The trajectory 

∆V’s were used in the ProbeMAASS1 Propulsion 

spreadsheet to estimate required propellant mass (fuel 

and oxidizer for the assumed bi-propellant, pressure 

regulated system) and propellant tank size. The 

pressurant mass and pressurant tank size were estimated 

in the spreadsheet from the propellant tank sizes and 

user input for pressurant molecular weight and system 

pressure. The actual propellant and pressurant tank 

masses were estimated by using the tank areal density 

(kg/m2

) for similar radius tanks16. Thruster masses 

were estimated from user input for number and size of 

thrusters, and a thruster mass trend curve6

. 

Miscellaneous propulsion system fitting mass was 

estimated as a user-specified mass fraction of the other 

propulsion system mass. Propulsion module dry mass 

was estimated by a simple linear scaling of a baseline 

chemical or SEP propulsion module, as appropriate. 

The propulsion system and propulsion module dry 

masses were then combined with other previously 

determined stackup component masses. Figure 2014, 

shows the resulting total launch growth wet mass vs. 

mission configuration for the ten different 

configurations listed above. ProbeMAASS1 allowed 

the user to show the mass-saving benefits of 

aerocapture missions to Titan for a variety of missions 

in a very short time. These particular results show that 

the aerocapture missions to Titan provide significant 

mass savings over all-chemical missions primarily due 

to the large propellant and propulsion module dry 

masses required for the all-propulsive ∆V maneuvers. 

CONCLUDING REMARKS

 The PROBECODES tool and its major component, 

ProbeMAASS1, are still in development. The goal is to 

have a near-seamless end-to-end tool for taking 

conceptual atmospheric entry probe designs, estimating 

system, subsystem, and component masses, and 

performing structural sizing optimizations on the probe 

aeroshells to produce a preliminary probe aeroshell 

structural design, all in a relatively short time. While 

not yet completed, great strides have been made 

towards completing this process, and portions of the 

tool have been successfully used to support mass 

sensitivity trade studies for the 2002 Titan Aerocapture 

Systems Analysis study. Enhancements will continue 

to be made as the tool develops. 

ACKNOWLEDGEMENTS 

 The author wishes to thank Dr. Mary Kae Lockwood 

of the NASA LaRC Vehicle Analysis Branch for her 

guidance in developing the tool architecture and tool 

focus. The author would also like to thank K. 

Chauncey Wu of the NASA LaRC Vehicle Analysis 

Branch for his inputs during the initial stages of the tool 

development, and Glenn A. Hrinda, also of the NASA 

LaRC Vehicle Analysis Branch, for his inputs for future 

tool enhancements. 

REFERENCES 

1. EDS PLM Solutions I-deas® software versions 

9 and higher, Electronic Data Systems 

Corporation, Plano, TX 

2. Microsoft® Excel User’s Guides 1 and 2, 

Microsoft Corporation, 1994. 

3. MSC/NASTRAN Quick Reference Guide, The 

MacNeal-Schwendler Corporation, 1992. 

4. Collier Research Corporation, HyperSizerTM 

Structural Sizing Software, Book 1: Tutorial 

Mission Launch 

Vehicle 

Gravity 

Assist 

Propulsion 

System 

Cruise 

Time (yrs.) 

Ventry 

(km/s) 

∆Vtrajectory

(km/s) 

1 Delta 4450 VVVGA Chemical 9.3 6.2 1.5 

2 Delta 4450 VGA SEP 5.7 6.5 8.3 

3 Delta 4450 EGA Chemical 7.5 5.3 1.6 

4 Delta 4450 EGA SEP 5.9 6.5 6.7 

5 Delta IV H VVVGA Chemical 9.3 6.2 1.5 

6 Delta IV H VGA SEP 5.7 6.5 6.9 

7 Delta IV H EGA Chemical 7.1 6.6 0.5 

8 Delta IV H EGA SEP 6.1 6.5 4.6 

9 Delta IV H VVVGA All chemical ~11.3 6.2 4.8 

10 Delta IV H VVVGA All chemical ~11.9 5.9 3.5 
119

& Applications, Second Edition, Collier 

Research Corporation, October 1998. 

5. Lockwood, M.K., et al, Titan Aeroshell 

Aerocapture Systems Analysis Review, Jet 

Propulsion Laboratory, Pasadena, CA, August 

29-30, 2002. 

6. Brown, C.D., Spacecraft Propulsion, AIAA 

Education Series, 1996. 

7. Staff of Research and Education Association, 

Handbook of Mathematical, Scientific, and 

Engineering Formulas, Tables, Functions, 

Graphs, Transforms, Research and Education 

Association, 1997. 

8. Bailey, R.W., Spacecraft Design, Titan 

Aeroshell Aerocapture Systems Analysis 

Review, Jet Propulsion Laboratory, Pasadena, 

CA, August 29-30, 2002. 

9. Laub, B., Thermal Protection (TPS), Titan 

Aeroshell Aerocapture Systems Analysis 

Review, Jet Propulsion Laboratory, Pasadena, 

CA, August 29-30, 2002. 

10. Congdon, W.M. and Curry, D.M., 2007 Titan 

Mission Analysis of ARA Ablators to Define 

Required Thickness and Weights, Analysis 

Results of July 11, 2002. 

11. Congdon, W.M., “Titan Ablator Thickness”, 

Analysis Results of August 23, 2002, e-mail 

from W.M. Congdon (Applied Research 

Associates) to M.K. Lockwood (NASA 

LaRC). 

12. Dec, J., “Genesis TPS for Titan Aerocapture”, 

Analysis Results of September 27, 2002, email from J. Dec (NASA LaRC) to M.K. 

Lockwood (NASA LaRC). 

13. Hrinda, G.A., Aeroshell Structure, Titan 

Aeroshell Aerocapture Systems Analysis 

Review, Jet Propulsion Laboratory, Pasadena, 

CA, August 29-30, 2002. 

14. Dyke, R.E., Mass Sensitivities, Titan Aeroshell 

Aerocapture Systems Analysis Review, Jet 

Propulsion Laboratory, Pasadena, CA, August 

29-30, 2002. 

15. Noca, M.A., Mission Analysis, Titan Aeroshell 

Aerocapture Systems Analysis Review, Jet 

Propulsion Laboratory, Pasadena, CA, August 

29-30, 2002. 

16. Sweetser, T.H., et al, “Titan Lander 

Conservative Science 01-06”, JPL Team-X 

Report, June 15, 2001. 
120

Figure 1. PROBECODES Entry System Mass Properties & Structural Design Tool Architecture 

Figure 2. ProbeMAASS1 Functional Features 

Figure 3. MPD/MPB Breakdown 

Payload 

Requirements (from 

Mission Profile)

Entry System 

Database

User Input 

Parameters Entry System Geometry

Solid Modeling 

(I-DEAS)

Geometric Data 

Transfer Link (GDTL)

Entry System Mass 

Properties Tool 

(ProbeMAASS1)

Mass Properties 

Output:

Lumped Masses

Aeroshell Mesh

Summary

Inertial, Aero, & 

Thermal Loads

Finite Element Modeling 

& Aeroshell Structural 

Sizing Materials Data

Aeroshell Structural Design & 

Probe Mass Estimate 

(Preliminary)

ProbeMAASS1

MPD: Mass 

Properties 

Database

MPB: 

Mass 

Properties 

Breakdown

MEM’s: 

Mass 

Estimation 

Methods

Aeroshell 

Automesher:

Surface Area 

Calculator

Lumpe d 

Mass 

Generator

PlateElement 

Aeroshell

Mesh 

Generator

I-DEAS 

Geometry 

Importer

Mission 

Overview

Total Mass 

Properties Geometry

Forebody

Aftbody

Entry 

Environment

Forebody

System

The rmal 

Protection

Structure

Separation 

Systems

The rmal 

Control

Miscellaneous

Aftbody

System

Thermal 

Protection

Structure

Separation 

Systems

Thermal 

Control

Miscellaneous

Deceleration 

Systems

Pilot 

Parachute

Main 

Parachute

Terminal 

Parachute

Airbag 

Systems

Payload

Structure

Science

Guidance/Navigation

Command & Data Handling

Communication

Power

Propulsion

The rmal Control

Harness/Cabling

Miscellaneous
121

Figure 4. ProbeMAASS1 Component Spreadsheets 

Figure 5. Sample MEM Curve from Historical Data

Total Aeroshell Forebody Structure Areal Density vs Loaded Surface Area

y = 0.022x + 4.4308

R2

 = 0.9627

y = 0.022x + 19.405

0

5

10

15

20

25

30

35

0 100 200 300 400 500 600 700 800

Loaded Surface Area: Design Entry g's x Forebody Surface Area (m2)

Structure Areal Density (kg/m 2

)

Polar 

Lander

Viking

Pathfinder

Huygens

Pioneer/Venus 2 Small

Galileo

Pioneer/Venus 2 

Large

Separated heatshield

Non-separated heatshield

Introduction

Usage Tips

Mission Definition

Payload

I-DEAS Geometry 

Import

Sphere-Cone 

Forebody

Sphere-Cone 

Aftbody

Parachute & 

Terminal Descent

Propulsion

Sphere-Cone Misc. 

Mass Items

Sphere-Cone 

Mass Summary

Sphere-Cone Lumped 

Mass Summary

Sphere-Cone Aeroshell 

Mesh Generator

Ellipsled Surface 

Area Calculator

Ellipsled Forebody

Ellipsled Backshell

Ellipsled Cruise Stage

Ellipsled Misc. Mass Items

Ellipsled Mass Summary

Ellipsled Aeroshell Mesh 

Generator

Celestial Body Database

Drag Coefficient Database

Sphere-Cone Database

Ellipsled Database

Aeroshell Deceleration Profile

Forebody TPS Parametrics

Sphere-Cone 

Branch

Ellipsled 

Branch

Supporting Databases
122

Figure 6. Typical Sphere-Cone Forebody and Aftbody Section Geometry 

Figure 7. Typical Sphere-Cone Aeroshell Plate Element Mesh from Automesher 

Figure 8. Typical Ellipsled Aeroshell Geometry

Figure 9. Typical Ellipsled Aeroshell Plate Element Mesh 

DA

hNC

rN θ1 LC

rS

DS

DIF

LS

θ4

Deff2

NOSE

CONE

SHOULDER

SHOULDER CONE

FOREBODY GEOMETRY

DB

Hr

REGION 1

REGION 2

REGION 3

REGION 4

REGION 5

REGION 6

r1bd

r1ht

r2ht

r3ht

r4ht

r5ht

r6ht

r1ang

r2ang

r3ang

r4ang

r5ang

r6ang

r1td, or r2bd

r2td, or r3bd

r3td, or r4bd

r4td, or r5bd

r5td, or r6bd

fdtd

hrad

Hemisphe re 

Cap scrad

scht

Sphe re Cap

Either as full 

backshells, or 

as part of an 

assembly

c Lc

Aeroshell Split

at Lines

ab

b

Reference 

plane

Possible TPS 

zone lines

γ

β

CL

nf1

θ

Configuration 10: Canted Upper Shell Separation Base above 

Reference Plane and Forward of Nose/Cylinder Interface

Cylinder Nose

Z Y

X Upper

Lower

X
123

Figure 10. Orbiter Wet Launch Mass Figure 11. Aerocapture System Mass vs. TPS 

 vs. TPS Material Material 

Figure 12. Aerocapture System Mass Fraction Figure 13. Forebody TPS Mass vs. TPS Material 

 vs. TPS Material 

 Figure 14. Aerocapture Ballistic Coefficient 

vs. TPS Material 

Titan Orbiter Wet Launch Mass vs Forebody TPS Material: 3.75 m 

Diameter, 6.5 km/s Entry, SLA B/S TPS 

900

950

1000

1050

1100

1150

1200

1250

SRAM-14

SRAM-17

SLA-561V

SRAM-20

SRAM-20 +1

SRAM-20 +2

PC-20

PC-20 +1

PC-20 +2

TUFROC

TUFROC +1

TUFROC +2

C-C

C-C +1

C-C +2

Forebody TPS Material

Orbiter Wet Launch Mass (kg)

50% Margin

30% Margin

Nominal

Thickness

Genesis Style +15%

Reference MEL with contingency = 1097.7 kg

*

* Baseline TPS

*

Titan Aerocapture System Mass vs Forebody TPS Material: 3.75 m 

Diameter, 6.5 km/s Entry, SLA B/S TPS

250

300

350

400

450

500

550

SRAM-14

SRAM-17

SLA-561V

SRAM-20

SRAM-20 +1

SRAM-20 +2

PC-20

PC-20 +1PC-20 +2

TUFROC

TUFROC +1

TUFROC +2

C-C

C-C +1

C-C +2

Forebody TPS Material

Aerocapture System Mass (kg)

50% Margin

30% Margin

Nominal

Thickness

Genesis Style + 15%

Reference MEL with contingency = 426.14 kg

Aerocapture dry mass = 426.14 kg

Titan insertion propellant mass = 88.9 kg

*

*

* Baseline TPS

Titan Aerocapture System Mass Fraction vs Forebody TPS Material: 3.75 

m Diameter, 6.5 km/s Entry, SLA B/S TPS

27%

29%

31%

33%

35%

37%

39%

41%

43%

45%

SRAM-14

SRAM-17

SLA-561V

SRAM-20

SRAM-20 +1

SRAM-20 +2

PC-20

PC-20 +1

PC-20 +2

TUFROC

TUFROC +1

TUFROC +2

C-C

C-C +1

C-C +2

Forebody TPS Material

Aerocapture System Mass Fraction

50% Margin

30% Margin

Nominal

Thickness

Genesis Style + 15%

Reference MEL with contingency =38.8% of Orbiter Launch Wet Mass

 = 41.5% of Aerocapture Mass

Based on 426.14 kg aerocapture system mass.

* Baseline TPS

*

*

Titan Aerocapture Forebody TPS Mass vs Material: 3.75 m Diameter, 6.5 

km/s Entry, SLA B/S TPS

0

50

100

150

200

250

300

SRAM-14

SRAM-17

SLA-561V

SRAM-20

SRAM-20 +1

SRAM-20 +2

PC-20

PC-20 +1

PC-20 +2

TUFROC

TUFROC +1

TUFROC +2

C-C

C-C +1

C-C +2

Forebody TPS Material

Forebody TPS Mass (kg)

50% Margin

30% Margin

Nominal Thickness

Reference MEL with contingency = 193.21 kg

Genesis Style +15%

* Baseline TPS

*

*

Titan Aerocapture Ballistic Coefficient vs Forebody TPS Material: 3.75 m 

Diameter, 6.5 km/s Entry, SLA B/S TPS, Cd = 1.44

53

55

57

59

61

63

65

67

69

71

73

SRAM-14

SRAM-17

SLA-561V

SRAM-20

SRAM-20 +1

SRAM-20 +2

PC-20

PC-20 +1

PC-20 +2

TUFROC

TUFROC +1

TUFROC +2

C-C

C-C +1

C-C +2

Forebody TPS Material

Ballistic Coefficient (kg/m2)

50% Margin

30% Margin

Nominal

Thickness

Genesis Style +15%

Reference MEL with contingency = 64.5 kg/m2

Based on 1026.11 kg aerocapture mass.

* Baseline TPS

*

*
124

Figure 15. Orbiter Launch Wet Mass Figure 16. Aerocapture System Mass vs. 

 vs. Diameter and TPS Diameter and TPS 

Figure 17. Aerocapture System Mass Fraction Figure 18. Forebody TPS Mass vs. Diameter and 

 vs. Diameter and TPS TPS 

Figure 19. Orbiter Ballistic Coefficient Figure 20. Total Launch Growth Wet Mass 

 Vs. Diameter and TPS vs. Mission Configuration 

Titan Orbiter Forebody TPS Mass vs Diameter: 6.5 km/s Entry, Max 

Growth, SLA B/S TPS

20

40

60

80

100

120

140

160

180

200

3 3.1875 3.375 3.5625 3.75

Aeroshell Diameter (m)

Forebody TPS Mass (kg)

TUFROC

PC-20

SRAM-20

SLA-561V

SRAM-17

SRAM-14

Reference MEL with contingency = 193.21 kg (3.75 m diameter)

Baseline

Titan Orbiter Launch Wet Mass vs Diameter: 6.5 km/s Entry, Max 

Growth, SLA B/S TPS

750

800

850

900

950

1000

1050

1100

3 3.1875 3.375 3.5625 3.75

Aeroshell Diameter (m)

Orbiter Launch Wet Mass (kg)

TUFROC

PC-20

SRAM-20

SLA-561V

SRAM-17

SRAM-14

Reference MEL with contingency = 1097.7 kg (3.75 m diameter)

Baseline

Titan Aerocapture System Mass vs Diameter: 6.5 km/s Entry, Max 

Growth, SLA B/S TPS

150

200

250

300

350

400

450

3 3.1875 3.375 3.5625 3.75

Aeroshell Diameter (m)

Aerocapture System Mass (kg)

TUFROC

PC-20

SRAM-20

SLA-561V

SRAM-17

SRAM-14

Reference MEL with contingency (3.75 m diameter):

426.14 kg aerocapture dry mass

88.9 kg Titan insertion propellant mass

Baseline

Titan Aerocapture System Mass Fraction vs Diameter: 6.5 km/s 

Entry, Max growth, SLA B/S TPS

24%

26%

28%

30%

32%

34%

36%

38%

40%

42%

3 3.1875 3.375 3.5625 3.75

Aeroshell Diameter (m)

Aerocapture System Mass Fraction

TUFROC

PC-20

SRAM-20

SLA-561V

SRAM-17

SRAM-14

Reference MEL with contingency (3.75 m diameter):

 = 38.8% of Orbiter Launch Wet Mass

 = 41.5% of Aerocapture Mass

Based on 426.14 kg aerocapture system mass.

Baseline

Titan Orbiter Ballistic Coefficient vs Diameter: 6.5 km/s Entry, Max 

Growth, SLA B/S TPS

52

57

62

67

72

77

82

87

3 3.1875 3.375 3.5625 3.75

Aeroshell Diameter (m)

Ballistic Coefficient (kg/m2)

TUFROC

PC-20

SRAM-20

SLA-561V

SRAM-17

SRAM-14

Reference MEL with contingency = 64.5 kg/m2

Based on 1026.11 kg aerocapture mass, 3.75 m 

diameter.

Baseline

Total Launch Growth Wet Mass vs Mission Configuration

0

1000

2000

3000

4000

5000

6000

7000

8000

9000

D 4450

VVVGA

Chem

D 4450

VGA SEP

D 4450

EGA

Chem

D 4450

EGA SEP

D IV H

VVVGA

Chem

D IV H

VGA SEP

D IV H EGA

Chem

D IV H EGA

SEP

D IV H

VVVGA

Isp 325

D IV H

VEEGA 

Isp 325

Mission Configuration/Launch Vehicle

Growth Masses (kg)

Lander

Orb/Lndr I/F

Orbiter dry

Aerocap Sys

Ejected Mass

PM/Orb I/F

Propellant

PM Dry

LV/PM I/F

All chemical payload growth mass @ Titan = 937 kg

Aerocapture

Constant aerocapture system growth mass = 515 kg

Constant payload growth mass @ Titan = 875 kg

All-chemical

D xxxx = Delta 

launch vehicle
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