John Riehl and Charles Trefny

Glenn Research Center, Cleveland, Ohio

Daniel Kosareo

ZIN Technologies, Inc., Brook Park, Ohio

SRM-Assisted Trajectory for the

GTX Reference Vehicle

NASA/TM—2002-211599

June 2002
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John Riehl and Charles Trefny

Glenn Research Center, Cleveland, Ohio

Daniel Kosareo

ZIN Technologies, Inc., Brook Park, Ohio

SRM-Assisted Trajectory for the

GTX Reference Vehicle

NASA/TM—2002-211599

June 2002

National Aeronautics and

Space Administration

Glenn Research Center

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NASA/TM—2002-211599 1 

SRM-ASSISTED TRAJECTORY FOR THE GTX REFERENCE VEHICLE 

John Riehl and Charles Trefny 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Daniel Kosareo 

ZIN Technologies, Inc. 

Brook Park, Ohio 44142 

ABSTRACT 

A goal of the GTX effort has been to demonstrate the feasibility of a single stage- to- orbit (SSTO) 

vehicle that delivers a small payload to low earth orbit. The small payload class was chosen in 

order to minimize the risk and cost of development of this revolutionary system. A preliminary 

design study by the GTX team has resulted in the current configuration that offers considerable 

promise for meeting the stated goal. The size and gross lift-off weight resulting from scaling the 

current design to closure however may be considered impractical for the small payload. In lieu of 

evolving the project’s reference vehicle to a large-payload class, this paper offers the alternative 

of using solid-rocket motors in order to close the vehicle at a practical scale. This approach offers 

a near-term, quasi-reusable system that easily evolves to reusable SSTO following subsequent 

development and optimization. 

This paper presents an overview of the impact of the addition of SRM’s to the GTX reference 

vehicle’s performance and trajectory. The overall methods of vehicle modeling and trajectory 

optimization will also be presented. A key element in the trajectory optimization is the use of the 

program OTIS 3.10 that provides rapid convergence and a great deal of flexibility to the user. This 

paper will also present the methods used to implement GTX requirements into OTIS modeling. 

INTRODUCTION

One goal of the GTX design effort is the generation of a 300 lb payload class SSTO conceptual 

vehicle of small size and weight. Small vehicles obviously require much less structure and 

therefore cost less to build and operate. The use of expendable external rockets can overcome 

shortcomings in propulsion or materials performance, and the vehicle concept can still evolve into 

a meaningful single stage to orbit vehicle. In effect, the GTX design team wanted to maintain a 

small vehicle, augment the thrust of the vehicle, and avoid deviating significantly from a single 

stage concept. An almost obvious design choice is the use of an existing solid rocket motor 

system that could be added to GTX without significantly disturbing critical flow paths, adding any 

more weight than absolutely necessary to accommodate the added thrust loads, and still allow 

the vehicle to steer. The GEM 46 solid rocket motors proved to be a good match for GTX. These 

motors are flight qualified having flown on the Delta III vehicle, have thrust vector controls, and 

appear to be readily available. 

In reference [1], Hack and Riehl present the methodology and models then used to simulate the 

GTX reference vehicle. Although that report addressed an earlier configuration of the vehicle, the 

methods, figures of merit, and problem solving approach remain the same. 
NASA/TM—2002-211599 2 

SYMBOLS 

A*

 Rocket Engine Throat Area 

Ac Vehicle Inlet Capture Area

Isp Specific Impulse 

M Mach number 

P0 Atmospheric Pressure 

P4 Duct Pressure 

Pc Rocket Engine Chamber Pressure 

q Dynamic Pressure 

T Net Thrust 

Wf Final Weight 

α Angle of attack (deg) 

γ Flight Path Angle w.r.t. horizontal (deg) 

ACRONYMS 

OTIS Optimal Trajectories by Implicit Simulation Computer Program 

RJ Ramjet Engine 

RAM Ramjet Engine 

SRM Solid Rocket Motor system 

SC Supersonic Combustion Ramjet Engine 

SCRAM Supersonic Combustion Ramjet Engine 

GTX PROPULSION 

In reference [2], Trefny describes the GTX RBCC propulsion in detail. Here we will briefly 

summarize the baseline GTX flight sequence and the various operating modes of the vehicle. 

The RBCC engine operates in four distinct modes. These modes, in the order that they occur in 

flight are: 

1. Combined-Cycle (Rocket/Ram) 

2. Ramjet (RJ) 

3. Scramjet (SJ) 

4. Rocket. 

The RBCC engine design is an integrated flow path that uses liquid oxygen and hydrogen 

propellants. An axisymmetric engine design is used to maximize the structural efficiency as well 

as to reduce some engine analysis uncertainties. Nozzles are integrated into the vehicle aft end 

in order to act as an expansion surface when operating at high speeds. An altitude compensating 

effect is provided by this arrangement at low speeds. 

The GTX launch vehicle employs three RBCC engine pods distributed 120° apart along the 

longitudinal axis and terminate at the rear of the vehicle. Figure 1 shows an isometric of the GTX 

with the GEM 46 SRM’s and figure 2 shows the same vehicle in a three-view drawing with 

dimensions in inches. 
NASA/TM—2002-211599 3 

ANALYSIS METHODOLOGY 

All of the flight design and trajectory optimization analyses reported herein were generated with 

version 3.10 of the computer program OTIS—Optimal Trajectories by Implicit Simulation

references [3 and 4]. OTIS implicit integration mode (also known as OTIS mode 4) was used 

exclusively. Implicit integration (or collocation) is used to simultaneously optimize and integrate 

the state differential equations. The states and control variables are represented by piecewise 

polynomials. The optimal control problem is then transcribed to a nonlinear programming problem 

via the implicit integration scheme. This is then solved by the nonlinear programming package 

SNOPT. To verify the converged trajectory, OTIS generates an explicitly integrated (4th order 

Runge-Kutta fixed step size) trajectory using the control history from the implicit solution. 

All vehicle modeling reported herein used only three degrees of freedom without any attempt to 

verify trim in the pitch plane. 

Flight path equations of motion were used for this problem just as in previous simulations without 

the SRM thrust augmentation. Additionally, flight path controls are used instead of Euler angle 

controls because of the aerodynamic flight involved in the problem. However, the only flight path 

control angle used in this study is the angle of attack, α, positive being nose up. The planar flight 

assumption yields zero bank and sideslip angles. 

FIGURES OF MERIT 

The primary figure of merit (FOM) for RBCC ascent trajectories is the weight of the vehicle in its 

final orbit, Wf. OTIS maximizes final weight, which in effect minimizes propellant. 

The overall ratio of oxidizer to fuel weights for the entire vehicle, O/F, provides a second figure of 

merit. This parameter is used to compare the effects of constraints on the split between the 

oxidizer and the fuel. 

The reader should note that the OTIS program does not directly compute O/F ratio. This is 

computed after the fact in post-processing software. 

CONTROLS 

Angle of attack, α, is the primary means of control for the RBCC trajectory optimization. This is an 

independent variable for all phases of powered flight after the vertical rise. During the vertical rise 

and during both atmospheric and vacuum coasts, α is set to zero to insure that drag is as small 

as possible. Angle of attack is limited to be within ±6° for modes 2 and 3 to insure sufficient inlet 

flow based on vehicle configuration. 

The chamber pressure, Pc

, in the rocket element during combined-cycle mode and full rocket 

mode is also a time varying control variable. Near take-off, the chamber pressure is higher to 

minimize the time in low speed flight. As velocity increases and the vehicle approaches the point 

where the ramjet mode can take over, the chamber pressure is allowed to decrease to 20% of the 

maximum. For this report, the maximum rocket chamber pressure is 2000 psi. In combined-cycle, 

the chamber pressure is constrained to always decrease. In rocket mode, the chamber pressure 

may also vary to down to 20% of the maximum to limit the acceleration of the vehicle. The 

baseline vehicle requires maximum chamber pressure to produce sufficient thrust for takeoff. For 

the thrust-augmented case, chamber pressure was optimized during the lift off of the vehicle to 

optimally match air breathing propulsion and the SRM propulsion. 

The duration of each propulsive mode is unspecified and open for optimization. As a practical 

matter, the mode durations are bounded by input to OTIS to insure good mathematical problem 
NASA/TM—2002-211599 4 

definition for the optimizer. Lastly, the initial azimuth after pitch-over, Ψ, is also open for 

optimization, but with 0  Ψ  

CONSTRAINTS 

The most significant constraint on the trajectory is the maximum dynamic pressure, q. This 

constraint affects the amount of heating and the pressure loads that the vehicle endures and 

sizes the thermal protection system, a significant vehicle design driver. The maximum dynamic 

pressure allowed over the entire trajectory is 1500 lb/ft2

. The minimum dynamic pressure for 

RJ/SJ mode is 500 lb/ft2

 in order to sustain combustion. From a trajectory standpoint, the vehicle 

cannot stay too low in the atmosphere while accelerating, as dynamic pressure becomes too 

high. 

Each engine mode is limited to operate within specified Mach number ranges. These ranges are 

shown in Table 1. Although the OTIS program enforces these ranges, actual mode transitions 

occur at Mach numbers that optimally provide the maximum benefit to maximizing final weight. 

A 10-atmosphere constraint on duct pressure prevents OTIS from selecting a flight regime that 

would induce internal vehicle pressures greater than can reasonably be accommodated by the 

vehicle structure. 

The vehicle is constrained to a maximum total acceleration of four times gravity (4 g’s) throughout 

ascent. This limitation tends to reduce the overall structural mass and prevents inducing loads on 

payloads beyond that typically found on existing launch systems. The vertical rise must be at 

least 500 feet to insure that any launch tower is cleared before controlled flight begins. Finally, the 

planar flight assumption yields zero bank and sideslip angles for the entire ascent. 

MODELS AND ASSUMPTIONS 

The trajectory optimization process requires a substantial amount of information about the 

propulsion modes and flight characteristics of the vehicle. This section discusses the types of 

data required by OTIS and how it is utilized. 

Propulsion Models 

SRM 

The thrust profile for the three GEM 46 motor is shown in figure 3. The Isp is 277.8 secs. and the 

total exit area is 22.3 ft2

. For simulation purposes in OTIS, this thrust versus time history is 

represented by a quintic chamfered spline. This representation provides the accuracy of linear 

interpolation while providing second order “smoothness.” The spline is called chamfered because 

it rounds off the “corner“ where linear interpolation segments meet. 

Mode 1 (Rocket-ramjet) 

The thrust and Isp in mode 1 are modeled as function of chamber pressure, atmospheric 

pressure, and Mach as follows: 

T = A*Pc fT1 M,

Pc

Po





  





  

Isp = fI1 M,

Pc

Po





  





  
NASA/TM—2002-211599 5 

where A*

 is thruster throat area in ft2, Pc is chamber pressure in lb/in2, P0 is atmospheric pressure, 

and the functions fT1 and fI1 represent mode 1 interpolating functions for the tabular data 

representations of GTX propulsion. OTIS interpolates this data using quintic or linear spline 

fits as appropriate in the dimensionality of the data. Total capture area for the GTX vehicle, Ac

, is 

138.2 ft2 and A*

 is .942 ft2. 

Modes 2 and 3 (Ramjet and Scramjet) 

The thrust and Isp for the ramjet and scramjet are calculated by similar relationships: 

T = qAc fTN ( ) M

Isp = fIN ( ) M

P4 = fIN (M)

These two engine operational modes have distinct tabular data, one for thrust coefficient, one for 

specific impulse, and one for the duct pressure (P4). These are represented above as fTN and fIN

where N=2 or 3. 

Mode 4 (Rocket) 

The rocket mode propulsion parameters are given as functions of engine throat area and 

chamber pressure. 

Tvacuum = A*Pc f

T4( ) Pc

Isp = fI4 Pc ( )

Here again fT4 and fI4 are mode specific interpolating functions for thrust and Isp. An atmospheric 

backpressure correction term is added to the rocket mode thrust value but has minimal impact to 

performance: 

T = Tvacuum − AexitPO

Rocket mode mixture ratio is also given as a tabular function that represents the impact of 

secondary flows within the engine that do not produce thrust. Hence the nominal mixture ratio of 

7/1 results in a mixture ratio of 6.77/1 for instance. 

Aerodynamic Model 

Similar to the propulsion model, tables for the coefficients of lift, CL, and drag, CD, are provided to 

OTIS. The APAS program reference [6] generated these coefficients for both the nominal 

configuration and the thrust-augmented configuration. In both cases, the data table interpolations 

are functions of angle of attack and Mach number. For the GTX configuration 10c used in this 

analysis, the reference area, Sref, is 188.84 ft2

. 

Earth Atmosphere and Gravitational Models 

The atmosphere model used for this study is the U.S. Standard Atmosphere, 1976. A quintic 

spline version of this atmosphere is used instead of direct evaluation in order to insure continuous 

first derivatives of the force terms affected by the atmosphere. No additional wind conditions are 

added to this model. Above 400,000 feet in altitude, the atmospheric model is set to a vacuum. In 

vacuum, the Mach number calculations use the speed of sound at standard sea level. Earth's 
NASA/TM—2002-211599 6 

Gravitational potential was assumed to be 1.40785E+16 ft3

/sec2 with an equatorial radius of 

6378.14 km or 20,925,656 ft. 

MAPPING PROBLEM TO ANALYSIS TOOL 

Within OTIS, a phase is defined as a segment of the trajectory with similar flight, aerodynamic, 

and propulsive characteristics. A number of trajectory nodes are placed within a phase. These 

are the points at which the implicit integration occurs. More nodes can increase precision but at 

the expense of longer the run-time. 

Table 2 lists controls and path constraints as given in Trefny reference [5] that are active within 

each operating mode. These constraints and bounds apply to the OTIS phases of both the 

baseline simulation and the thrust-augmented simulation. These two simulations differ in their 

structure so that the thrust-augmented simulation ends precisely at SRM depletion, the use of 

SRM thrust is stopped and in a subsequent OTIS phase, the burned out SRM boosters are 

jettisoned. From thereon to orbit insertion, the simulations are identical. 

MISSION AND FLIGHT SEQUENCE OF THE THRUST AUGMENTED VEHICLE 

The baseline mission utilized for this study is the planar ascent of the GTX vehicle to a 220 Nmi 

orbit at 28.5° inclination. 

The SRM’s ignition begins the flight of the thrust-augmented GTX. The RBCC’s first operational 

mode, rocket-ramjet, is also operating and generating thrust. The chamber pressure is not at its 

maximum value as in the baseline GTX, but at a value determined that maximizes on-orbit (final) 

weight. Operating in this mode, the vehicle begins with a vertical rise followed by a pitch-over. 

The vehicle accelerates through the transonic region until reaching a speed where ramjet 

combustion can be sustained. The rocket-ramjet mode is not nearly as efficient as the ramjet 

mode. So one objective of the trajectory optimization process is to determine the best flight 

conditions for where the handoff between Combined-Cycle and ramjet occurs. The GEM 46 stops 

producing significant thrust at 77 seconds. At 80 seconds into the flight the SRM’s are presumed 

to have burned out completely. They are jettisoned and the GTX core vehicle continues its flight. 

The ascent continues in ramjet mode as the vehicle accelerates, eventually switching to 

supersonic combustion ramjet mode. These middle two operational modes use only atmospheric 

oxygen, further reducing the on-board oxygen storage requirements. As such, they are the most 

efficient segments of the ascent. OTIS maximizes the use of these stages while satisfying 

associated path constraints. 

Finally, a pure rocket mode, using stored LOX, takes over and accelerates the vehicle through 

the atmosphere. The rocket is throttled to achieve the acceleration constraints on the trajectory. 

This, in turn, decreases the weight of the structural subsystem. After orbital insertion velocity is 

achieved, the vehicle coasts to just below the final orbit altitude, where a circularization burn in 

rocket mode occurs, ending the ascent simulation. 

PERFORMANCE COMPARISON OF BASELINE MISSION AND THE THRUST 

AUGMENTED MISSION 

The gross liftoff weight of the baseline vehicle is 236,000 pounds while the addition of the three 

SRM’s increases this weight to 360,770 pounds for the thrust-augmented vehicle. The extra 

weight comes from the three loaded GEM 46 SRM’s exclusively. 
NASA/TM—2002-211599 7 

Tables 3 and 4 present summaries of key trajectory events and performance for the baseline and 

the thrust-augmented cases, respectively. In each table the baseline value is presented above 

the thrust-augmented value. In table 3 one can see that the thrust-augmented case rises faster 

and transitions into the first three propulsion modes sooner than the baseline case. This is as 

expected, but these faster transitions come at the expense of altitude attained relative to the 

baseline vehicle. It remains in mode 3 longer than the baseline to take complete advantage of the 

more efficient air breathing modes longer than the baseline even though the transition to mode 4 

occurs at the same Mach number. The first rocket burn is some seven seconds longer because 

the thrust-augmented vehicle is heavier than the reference at the start of the burn and lower in 

altitude. Atmospheric coast to the second rocket burn (orbital injection burn) takes longer 

because OTIS has found a different transfer orbit (apogee altitude = 224.3 x perigee altitude = 

–92.9 Nmi. nominally vs. 228.5 x –46.2 Nmi. for the thrust-augmented case). The final altitudes of 

the nominal and thrust-augmented cases differ slightly because OTIS converges to a specified 

tolerance on final apogee and perigee altitude (220 ± 0.5 Nmi.) over a non-spherical earth. 

In table 4 one observes that the thrust-augmented vehicle uses about the same amount of 

hydrogen to get on orbit and considerably less oxygen. The rocket only mode (mode 4) requires 

more oxygen and hydrogen than the baseline simply because the vehicle weighs more at the 

start of the mode. 

Figures 4 to 13 show comparisons between the baseline and the thrust-augmented vehicles’ key 

trajectory parameters. Note that the angle of attack history exhibits a series of rapid oscillations in 

mode 3 that persist even after an integral compensation feature was employed to remove them. 

Also limiting the rate of change in angle of attack produced less final weight. One can only 

conclude that the resolution of the propulsion and aerodynamic data and implicit integration 

scheme have coupled in such a way as to produce these oscillations in the vicinity of a truly 

optimal solution. With this sole exception there is little that is unexpected in the ascent of the 

thrust-augmented vehicle relative to the nominal. Vehicle constraints in dynamic pressure and 

total acceleration are met while always transitioning from engine mode to engine mode at the 

appropriate Mach number. 

One can reasonably conclude that the thrust-augmented GTX will supply the added weight on 

orbit that is desired. The final weight represents 24.5% of the initial GTX weight of 236,000 lbs. as 

opposed to the nominal vehicles 20.7%. The extra margin could absorb future weight and 

performance shortfalls. 

SUMMARY 

Thrust augmentation using the GEM 46 provides the additional weight on orbit that the GTX 

design team sought without significantly compromising the intent of single stage to orbit 

transportation. These SRM’s are flight qualified, available, and steerable. There installation on the 

GTX conceptual vehicle appears to be straightforward. 

REFERENCES 

1. Hack, Kurt J., and Riehl, John P., ”Trajectory Development and Optimization of an RBCCbased Launch Vehicle,” Paper AAS 99–347AAS/AIAA Astrodynamics Specialist 

Conference, Girdwood, Alaska, 16–19 August 1999. 

2. Trefny, C.J., “GTX Program Overview,” Proceedings of the JANNAF 25th Airbreathing 

Propulsion Subcommittee Meeting, November 13–17, 2000, Monterey, California. 
NASA/TM—2002-211599 8 

3. Hargraves, C.R., and Paris, S.W., Direct Trajectory Optimization Using Nonlinear 

Programming and Collocation, Journal of Guidance, Control, and Dynamics, Vol. 10, No. 4, 

July–August 1987. 

4. Hargraves, C.R., and Paris, S.W., Optimal Trajectories by Implicit Simulation (OTIS 3.0) 

Volume I – Formulation Manual, Boeing Defense and Space Group, June 1996. 

5. Trefny, C.J., Trajectory Optimization Task Description. GRC Internal Memorandum, March 

7, 2001. 

6. Sova, G., and Divan, P., Aerodynamic Preliminary Analysis System II, Parts I and II, 

Rockwell International Corp. 

Table 1. RBCC ENGINE MODE CONSTRAINTS 

Engine Mode Minimum 

Mach 

Maximum 

Mach 

Minimum Duct 

Pressure (Atm.) 

Maximum Duct 

Pressure (Atm.) 

Combined-Cycle – 3 N.A. N.A. 

Ramjet 2.5 4 0.5 10 

Scramjet 5 15 0.5 10 

Rocket 4 N.A. N.A. N.A. 

Table 2. SIMULATION CONTROLS AND PATH CONSTRAINTS 

 Mode 1 Mode 2 Mode 3 Mode 4 

 Min Max Min Max Min Max Min Max 

Freestream dynamic pressure, Q0 (psfa) 500 1500 500 1500 500 1500 500 1500 

Angle-of-attack, α (degrees) –6 6 –6 6 

Thruster chamber pressure, Pc (psia) 400 2000 400 2000 

Diffuser exit pressure, P4 (atm) 1/2 10 1/2 10 

Total acceleration (g's) 4 4 4 4 
NASA/TM—2002-211599 9 

Table 3. KEY MISSION EVENTS FOR THE BASELINE AND THRUST AUGMENTED VEHICLES 

Flight Time Altitude 

Relative 

Velocity Mach Weight 

Phase (sec) (ft) (Nmi) (ft/sec) (lb) 

Liftoff 0 

0 

0 

0 

0 

0 

0 

0 

0 

0 

236,000 

360,770 

End of 

Vertical 

Rise 

6.1 

5.35 

500 

500 

0.08 

0.08 

163 

191.4 

0.15 

0.17 

229,027 

347,003 

Mode 1–2 65.1 

53.5 

43,015 

43,015.3 

7.1 

7.1 

2414.5 

2414.5 

2.5 

2.5 

174,017 

240,633 

End SRM N.A. 

80 

N.A. 

62,696.2 

N.A. 

10.3 

N.A. 

3801.5 

N.A. 

3.9 

N.A. 

217,134 

then 

203,700 

Mode 2–3 149.6 

114.3 

78,277 

72,242 

12.9 

11.9 

5,374 

4,868.6 

5.5 

5 

167,455 

201011 

Mode 3–4 613.5 

617.3 

122,443 

112,838 

20.2 

18.6 

11,183 

10,821 

10.9 

10.7 

133,314 

169,659 

End of 1st 

rocket burn 

720.3 

731.8 

228,321.5 

217,737.4 

37.6 

35.8 

24,568.5 

24,673.2 

25.1 

24.7 

50,996.5 

59,937.5 

Start 2nd 

rocket burn 

2,201.8 

2,421.8 

1,340,829.8 

1,322,160.8 

220.7 

221.2 

23,160.4 

23,245.3 

20.7 

20.8 

50,996.5 

59,937.5 

End 2,209.8 

2,429.9 

1,340,940.2 

1,344,223.2 

220.7 

221.2 

23,731.8 

23,729.4 

21.3 

21.3 

48,827.9 

57,769.1 

Table 4. PROPELLANT BREAKDOWN FOR THE BASELINE AND 

THRUST AUGMENTED VEHICLES 

Engine Mode Oxidizer Weight 

(lb.) 

Fuel Weight (lb.) Total Propellant 

(lb.) 

O/F Ratio 

Mode 1 52,038 

24,693 

9,533 

5,415 

61,571 

30,108 

6.34 (aver.) 

5.90 (aver.) 

Modes 2 and 3 0 

0 

34,125 

36,104 

34,125 

36,104 

N.A. 

N.A. 

Mode 4 79,003 

97,207 

11,671 

14,360 

90,674 

111,567 

6.77 

6.77 

Total 131,041 

121,900 

55,329 

55,879 

186,370 

177,779 

2.37 

2.18 
NASA/TM—2002-211599 10 

Figure 1. Isometric view of GTX with SRM added 

Figure 2. Three-View of GTX 10C with GEM 46 SRM's 
NASA/TM—2002-211599 11 

Figure 3. GEM 46 thrust profile 

Figure 4. Angle of Attack vs. Mach 
NASA/TM—2002-211599 12 

Figure 5. Flight Path Angle vs. Mach 

Figure 6. Weight vs. Flight Time 
NASA/TM—2002-211599 13 

Figure 7. Altitude vs. Down Range Distance 

Figure 8. Altitude vs. Mach 
NASA/TM—2002-211599 14 

Figure 9. Chamber Pressure vs. Mach 

Figure 10. Acceleration vs. Mach 
NASA/TM—2002-211599 15 

Figure 11. Dynamic Pressure vs. Mach 

Figure 12. Total Vehicle Thrust vs. Mach 
NASA/TM—2002-211599 16 

Figure 13. Effective Isp vs. Mach 
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10. SPONSORING/MONITORING

 AGENCY REPORT NUMBER

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546–0001

Available electronically at http://gltrs.grc.nasa.gov/GLTRS

June 2002

NASA TM—2002-211599

E–13386

WU–708–90–63–00

22

SRM-Assisted Trajectory for the GTX Reference Vehicle

John Riehl, Charles Trefny, and Daniel Kosareo

Airbreathing propulsion; Launch vehicles; Hypersonic propulsion; Trajectory analysis

and optimization

Unclassified -Unlimited

Subject Categories: 15 and 07 Distribution: Nonstandard

Prepared for the Combustion, Airbreathing Propulsion, Propulsion Systems Hazards, and Modelling and Simulation

Subcommittees Joint Meeting sponsored by the Joint Army-Navy-NASA-Air Force, Destin, Florida, April 8–12, 2002.

John Riehl and Charles Trefny, NASA Glenn Research Center; Daniel Kosareo, ZIN Technologies, Inc., Brook Park,

Ohio 44142. Responsible person, John Riehl, organization code 7820, 216–977–7061.

A goal of the GTX effort has been to demonstrate the feasibility of a single stage-to-orbit (SSTO) vehicle that delivers a

small payload to low earth orbit. The small payload class was chosen in order to minimize the risk and cost of development of this revolutionary system. A preliminary design study by the GTX team has resulted in the current configuration

that offers considerable promise for meeting the stated goal. The size and gross lift-off weight resulting from scaling the

current design to closure however may be considered impractical for the small payload. In lieu of evolving the project’s

reference vehicle to a large-payload class, this paper offers the alternative of using solid-rocket motors in order to close

the vehicle at a practical scale. This approach offers a near-term, quasi-reusable system that easily evolves to reusable

SSTO following subsequent development and optimization. This paper presents an overview of the impact of the addition

of SRM’s to the GTX reference vehicle’s performance and trajectory. The overall methods of vehicle modeling and

trajectory optimization will also be presented. A key element in the trajectory optimization is the use of the program

OTIS 3.10 that provides rapid convergence and a great deal of flexibility to the user. This paper will also present the

methods used to implement GTX requirements into OTIS modeling.