NASA Technical Memorandum 58264

(B&SA-TM-5%26_) ]:/_5_GN 5_[UD_ O_ /iN N85-21262

. II_T_GB..21_E AE_OEBAKING OBEI_IAI _RJlGSFER

V_HICL_. (NASa) _,_ p BC AOS/M_ J.Ul CSCI. 22B

Oncla_

• G3/20 15147

Design Study of an Integrated

Aerobraking Orbital

Transfer Vehicle

March 1985

NationalAeronautics and

SpaceAdministration
NASA Technical Memorandum 58264

#

I

Design Study of an Integrated Aerobraking i

Orbital Transfer Vehicle

!

Carl D. Scott, Barney B. Roberts, Kornel Nagy, Peter Taylor,

Joe D. Gamble, Christopher J. Ceremeli, Kenneth R. Kroll,

Chien P. Li, and Robert C. Ried

March 1985

National Aeronautics and Space Administration

Lyndon B. Johnson Space Center

Houston, Texas 77058
"I

'I

/

CONTENTS

Section Page

ABSTRACT...........................................................1

INTRODUCTION.......................................................1

AEROASSIST BACKGROUND..............................................2

VEHICLE CONCEPT AND GROUND RULES................................... 6

GEOMETRICAL CONCEPT DESCRIPTION.................................... 7

AERODYNAMICS.......................................................7

VEHICLE CONCEPT....................................................7

PROPULSION.........................................................8

VEHICLE STRUCTURE CONCEPT..........................................9

STRUCTURAL MASS ESTIMATE...........................................I0

AEROBRAKING TRAJECTORY.............................................10

AERODYNAMIC HEATING ESTIMATES......................................II

THERMAL PROTECTION.................................................12

TOTAL VEHICLE MASS.................................................12

PERFORMANCE........................................................13

TECHNOLOGY IMPROVEMENTS............................................15

CONCLUSIONS........................................................t5

REFERENCES.........................................................32

TABLES

Tab]e Page

1 JSC AOTV CONFIGURATION MASS SUMMARY (GEO RETURN).............16

I'!¢J_.CEI)I}_:_ P.%C3F.BF.A_IK NOT P,_l'._fI_._

ill
FIGURES

Figure Page

1 Payload-to-stagemass ratio of all propulsive and aerobraked

stages....................................................17

2 Improvement in GEO delivery payload compared with aerobrake

mass......................................................18

3 Geometric construction of klunted, raked-off cone............ ]9

4 Integrated AOTV design in aerobraking attitude...............20

5 Newtonian aerodynamic coefficients of blunted, raked-off

cone......................................................21

6 Integrated AOTV concept with 12-m (40-ft) diameter heat

shield....................................................22

7 Side view of half-model AOTV heat shield structure and

bulkheads.................................................23

8 Front view of half-model of AOTV heat shield structure and

bulkheads.................................................24

9 Altitude time history during AOTV aerobraking................25

10 Velocity time history during AOTV aerobraking................26

11 AOTV atmospheric flight trajectory characteristics...........27

12 AOTV atmospheric flight environment..........................28

13 Heat flux and pressure distributions on aerobrake

referenced to stagnation point value.......................29

14 Computational body with Mach number contours in shock

layer......................................................30

15 Fractional increase of GEO delivery mass for aerobraked

stage over an all propulsive stage.........................31

iv

,%
ACRONYMS

AOTV Aerobraking orbital transfer vehicle

EOM End-of-mission

GEO Geosynchronous Earth _,tbit

HRSI High temperature reuseable surface insulation

LEO Low Earth orbit

OTV Orbital transfer vehicle

RCS Reaction control system

RTV Room temperature ,u_canizingsilicc _ rubber

TieS Thermal protection system

SYMBOLS

A reference area

a dimensionless ratio, AV/Isp go

CD drag coefficient

CL lift coefficient

Cm pitching moment coefficient

Cn yawing moment

Isp specific impulse of rocket

L/D lift-to-drag ratio

M Mach number

m mass of vehicle at entry

q heat flux

RN nose radius at stagnation point

t effective skin thin thickness

V velocity

V

m ml l_#
SYMBOLS CONTINUED

WAB weight penalty for aerodynamic and TPS surfaces

WGE0 mass of payload delivered to GEO

WGEOAP payload mass to GEO by all propulsive stage

WLE0 total departure mass from LEO to GEO

Wp full propellant mas_ capacity

WS stage dry mass

WSAB aerobraked stage mass

WSAP propulsive stage mass

Wf weight at end of burn

Wi weight at start of burn

angle of attack

side slip angle

rake angle

_ bluntness ellipticity j b

-I

_ mass fraction

-I

p density

g cone angle

SUBSCRIPTS

i

AB, B aerobrake

AP, P all propulsive

!

_ freestream condition

vl
ABSTRACT

An aerobraking orbital transfer vehicle (AOTV) concept, which has _n aerobrake structure that is integrated with _he propulsion stage, is discussed.

The conceptual vehicle is to be assembled in space and is space-based. The

paper discusses the advantages of aeroassist over an all propulsive vehicle

and indicates that the vehicle considered is very competitive with inflatable and deployable concepts from mass and performance aspects. The aerobrake geometry is an ellipsoidaIly blunted, raked-off, elliptical wideangle cone with a toroidal skirt. Propellant tanks, engines, and subsystems are integrated into a closed, isogrid aerobrake structure which provides rigidity. The vehicle has two side-firing, gimbaled RL-IO type

engines connected by struts to a web bulkhead thrust structure.

The AOTV has 38 000 kg (84 000 Ib) of useable propellant loaded into tanks

and is sized to be carried in the Shuttle payload bay. One of the liquid

hydrogen tanks is sized for just the descent to reduce boiloff. The trajectory during aerobraking is determined from an adaptive guidance loqic,

and the heating is determined from engineering correlations as well as 3-D

Navier-Stokes solutions. A thermal protection system (TPS) consisting of

high temperature surface insulation bonded to the skin of the aerobrake is

sized for the expected aeroheating. The aerobrake structure and TPS are

found to be approximately 17 percent of the total stage dry mass. The AOTV

is capable of placing a 13 500 kg (29 800 lb) payload into geosynchronous

Earth orbit (GEO) or carrying a LEO-GEO-LEO round-trip payload of 7100 kg

(15 700 Ib). A two-stage version considered for lunar missions resultc in

a lunar surface delivery capability of 18 800 kg (41 500 Ib) or a roundtrip capability of 6800 kg (15 000 Ib) with an additional 3860 kg (8500 lb)

deliver-only capability. The dry mass of the stage is approximately 4430

kg (9760 Ib) and the re-entry mass with a 6800 kg (15 000 Ib) payload is

13 200 kg (29 000 Ib).

INTRODUCTION

As the Space Transportation System becomes an operational reality and the

Nation addresses the challenge _f permanent nanned presence in space, it is

appropriate to consider practical approaches toward the transport of systems, materials, and man beyond the low Earth orbit (LEO) region of the

Space Station. Given the economy of the reuseable Space Shuttle Orbiter,

the energy requirements for leaving the Earth to achieve LEO are dear and

J not to be squandered. This, in conjunction with a Space Station in LEO,

_ drives one toward consideration of space-based reuseable vehicles designed

i to take advantage of a more specialized application. The vehicle should be

i versatile enough to meet mission requirements, but not carry the design

L burdens required for returning to the surface of the Earth or the energy

expense of reboost to LEO. The achievement of this efficiency is not without investment. The design and development of space-based transportation

systems entails challenge for the assembly, servicing, maintenance, check-

!_ out, loading of propellant and cargo, and all aspects of vehicle turnaround !

in an Earth orbital environment.

!

h
The second major energy efficiency to be considered is the use of atmospheric braking for the dissipation of relative energy upon return to LEO.

,_ This maneuver avoids using propellant masses which would otherwise be carried for a full mission, but requires the mass burdens of an aerobraking

thermal protection system (TPS) and the design constraints of an acceptable

aerodynamic configuration. The Mercury, Gemini, Apollo, and Space Shuttle

Orbiter have used aerobraking to great advantage for return to Earth with

the latter system having achieved this with a reuseable TPS system. The

"_ reuseable TPS feature is of great advantage for a space-based system, but

Z more difficult as the relative velocity and the aerodynamic heating increase. Aerobraking to LEO is also significantly different from atmospheric entry achieved by previous manned spacecraft in that capture must be

avoided. Thus, an aerobraking orbital transfer vehicle (AOTV) must fly between the limits of the atmosphere and a capture boundary as opposed to,

, for example, the Apollo vehicles which flew between a capture boundary and !

severe deceleration limits. Like the Space Shuttle Orbiter, an AOTV flight

through the atmosphere must avoid excessive heating and must be consistent

with the TPS capability and the aerodynamic performance requirements for

achieving the maneuver.

The purpose of this paper is to project the potential characteristics of an

AOTV by consideration of a particular or "point" design based on reasonable, but somewhat arbitrary guidelines. The attempt is to extrapolate

slightly into the future based on the historic trends of both systems and

technology. The value of this is fourfold. First, the concept of what a

space-based AOTV might be like should help in consideration of the design

and development of a Space Station, at which such a vehicle may be based.

Second, the development of a propulsion system for an AOTV requires a concept of what the overall system might entail. Third, a base of vehicle

characteristics predicted on a somewhat conservative approach of extrapolation from the past is needed to provide perspective on current and future

innovative ideas and approaches. Finally, identification of potential system characteristics should help provide guidance for planning programs in

related research and technology development.

In perspective, the design process does not provide unique solutions and

seldom provides optimum solutions. Given the same guidelines as selected

here, there are a large number of approaches toward achieving the same end.

On the other hand, it is felt that the results presented here provide a

reasonable characterizationof an AOTV subject to the guidelines used and

any great leaps in technologicaldevelopment.

AEROASSIST BACKGROUND

Several concepts have been proposed for NASA's Orbital Transfer Vehicle

iOTV), and performance trade studies have been performed to compare these

concepts. This background section will address those trade studies that

were accomplished to establish the rationale for using aerobraking rather

than propulsive braking as well as why the integrated structural concept

was proposed for this study. The first study was a comparison of the ratio

oF initial mass in LEO to the payload mass carried to GEO. Since the cost

of each unit of mass into LEO is fixed at $2860 per kilogram ($1300 per

pound) at current (1984) Shuttle delivery rates, this mass ratio gives a

2
II

direct cost indicator for each proposed configuration. The higher the mass

ratio (defined as WLEO/WGEo), the higher the cost for delivery to GEO.

These mass ratios can be determined analytically as a function of the stage

propellant mass fraction X = Wp/(WS + Wp). This independent parameter,

mass fraction, was chosen because historical design data is available that

show, to a first order approximation,that the dry weight of an orbit

transfer stage is prcportional to the full propellant capacity of the stage

Oh"

WS : (I- l) Wp

i where

i k = mass fraction

I

Wp = full propellant mass capacity

WS = stage dry mass

The aerobraked stage can be characterized in much the same way for convenience and comparison purposes. The scaling law chosen for the aerobraked

stage is

WSAP WAB

= and _ =

WSAB 1 - _ B

B WSAP + WAB

where

WSAB = aerobraked stage mass

WSAP = mass of an all propulsive stage with equal propellant capacity as

the aerobraked stage

WAB = weight penalty for aerodynamic and TPS surfaces

B = aerobrake weight penalty fraction

Then, with the above relationships, the rocket equation

Wi

V = golspln

where

3
V = velocity change

go = units conversion constant 32.2 Ibm-ft/(Ibf-sec2) in U.S.

Engineering System

I : specific impulse of the rocket

sp

Wi = weight at start of burn

Wf = weight at end of burn

will allow solution of the functional relationships needed for the weight

ratio WLEO/WGEO. These solutions are

-1 + e'(al+a2)

WLEO/WGE0 = for the all propulsive stage

-I + e"2(a + a )

i 2

and

-i + e'a2

WLEo/WGE0 : for the aerobraked stage

-I + e'(al+2a2 )

where WGE0 is the mass of payload delivered to GEO, WLE0 is total departure

mass, and

AVI A V2 _V3

aI : a2 = a3 = where

Isp go Isp go Isp go

AVI : 2440 m/sec (8000 ft/sec), GEOtransfer orbit insert,on

' _V2 = 1830 m/sec (6000 ft/sec), GEO circularization

AV3 = go m/see (300 ft/sec), required velocity change necessary to circularize the vehicle's orFit after the aerobraking maneuver

Figure 1 shows the mass ratio for both an aerobraked stage and an all propulsive stage (with Isp = 4510 m/s or 460 Ibf-sec/Ibm) plotted as a function of stage mass fraction _ . The asymptotes 0.7411 and 0.8492 are the

stage mass fractions at which the stage has zero delivery capability to

GEO. First, note that for a given mass fraction, the aerobraked stage substantially outperforms the all propulsive stage. At a mass fraction of

0.86 (equivalent to a Centaur vehicle) the mass ratio for an aerobraked

stage is 4.5, whereas that for a propulsive stage is 22. However, this is

an invalid comparison since an aerobraked stage will be heavier than an a11

propulsive stage because of the mass required for the aerobrake and thermal

protection systems.

4
Prior to the design concept described in this report, the aerobraked vehicles proposed to the agency were more like all propulsive stages with an

aerobrake subsystem "bolted-on",much like an afterthought. The straight

lines connecting the two exponential curves actually link the design points

for the two different stages with the mass ratio of the aerobrake suFs__-.

tems indicated on the lines. This ratio is computed as a fraction c' ,.._L

total stage dry mass. For example, an all propulsive stage of ma_ frac-.

tion 0.86 with "bolted-on" aerobrake subsystems that weigh as much as _C

percent of the total vehicle dry mass, will equate to the aerobraked stage

joined by the line labeled " _ B = 0.4." Note that for the case disc1_sRed

the 40 percent aerobrake is not very efficient. A 30 percent aerobrake is

more efficient, but not a windfall benefit as mass fraction increases. Another way to measure the benefit of the aerobrake is shown in figure 2

where the increase in payload to GEr 'or an aerobraked stage is plotted as

a function of the brake mass ratio and the mass fraction of the all propulsive stage to which it is "bolted." This functional relationship is determined in much the same manner as the weight ratios in figure 1. The aerobraked OTV's weight penalty is defined differently to facilitate the calculation. However, the concept of aerosurface penalty is sti]l similar.

Now

WSAB - WSAP =

AB

WSAB

where

= weight penalty factor AB

Assuming that the two competing OTV's have the same propellant capacity,

use the rocket equation to yield

2(a +a2) eal_2a2+a3-1

WGE0 [e I -I] I -_ AB

2(a1+a2)

WGEOAP _ AP e -I

i - X AP

Here WGEOAP is the GEO delivery capability of the all propulsive stage,

AWGE 0 is the difference between the delivery capability of the aerobraked

stage and the all propulsive stage, and _ AP is the propellant mass

fraction of the all-propulsive stage.

Again, aerobrakes that are 30 to 40 percent of total dry mass do not accrue

much in increased delivery capability. However, previous OTV design studies from aerospace contractors show aerobrake masses to be In this range.

Therefore, one must conclude that if aerobraked stages cannot be constructed in a more efficient manner, then they are probably not worth the

5
" | 1

; ddditiondl c,)mpIexityin vehic]e design or opPration. It was this conclusion that prompted consideration of possible (:onceptsof integrated structures wherein structural support of tanks, engines, thermal protection systems, and other subsystems is shared in order t,)reduce aerobrake weight

penalty. The concept described herein achieve that goal.

VEHICLE CONCEDF AND GROUND RULES

The AOTV concept is an ell'psoidally blunted, raked-off elliptical cone and

is an extension of the concept evaluated from an aerothermal and thermal

protection viewpoint in reference I.

_any concl)ts of orbital transfer vehicles have been considered over the

last few years (refs. 2-7). All-propulsive vehicles use rocket thrust to

produce the necessary velocity decrease to achieve a LEO _fter return from

GEO. Aeroassisted vehicles use the drag of the atmosphere to accomplish

the majority of the velocity decrease. There have been a number of concepts investigated in the past including the ballute; the fixed drag brake

offering no lift; lifting brakes, including both high and low lift concepts; and aeromaneuvering vehicles that have high lift. Most of these

concepts are envisioned as having an aerobrake device attached to an existing cylindrical propulsion module that would fit in tne payload bay of the

Space Shuttle Orbiter. This study considers a concept that departs from

that requirement in that it is an integrated syst_n that would be assembled

in space. The aerobrake structure and the components of the propulsion and

other systems are integrated to obtain structural efficiency and therefore

s,nallermass.

The AOTV concept assumes certain ground rules and features. The vehicle is

designed to be space-based; that is, it may be assembled in orbit at a

space station and is not required to be capable of returning to Earth in

the Shuttle payload bay. Components will be brought to LEO from the ground

_ in the Shuttle payload bay for assembly. The vehicle is propelled by liquid hydrogen/oxygenengines. The total useable propellant mass assumed for

tnis study is 38 000 kg (£4 000 lb), which defines the basic capability of

the vehicle. This particu]ar amount of propellant was chosen for comparison with other concepts studied in the past having the same total

propellant.

The present concept assumes the delivery of payloads to GEO and the return

of payloads including a 6800 kg (15 000 lb) manned capsule to LEO. The vehicle must be reuseable, reliable and require a minimum of maintenance or

refurbishment. Operationally, it is envisioned that the vehicle will make

a plane change from a 28.5 inclination orbit to an equatorial orbit and

back by propulsive impulse. The majority of the velocity derrease in returning from GEO to LEO will be accomplished by aerobraking. The need for

lift is relegated to control of the vehicle in the atmosphere and not for

significant plane changes.

Because of the possibility of a manned return, a single atmospheric pass is

baselined

6
An aluminum structure is baselined, although composite materials are known

to be more mass-effective and, therefore, may be preferable in the ultimate

vehicle.

GEOMETRICAL CONCEPT DESCRIPTION

The geometical shape of the aerobrake structure was ch_sen to be a blunted

wide angle cone that serves as part of the structure of _he stage (figs. 3

and 4). The particular conR is an e11iptical cone (8=60 in the x-y plane)

raked-off at an angle of 73_. The wide-angle cone was selected beca_sc it

has a flatter shape for surface area efficiency and it provides lower stag-

;_ationpoint heating than sharper cones. The base of the cone (in the rake

plane) is c_rcular, t_,usdefining, along with the rake angle and cone angle, the con3 plllpt_city. A circular base is not required, but is considered here for packaging reasons.

An ellipsiodal bluntness was chosen which is tangent to the e11iptic cone

in a planar intersection perpendicular to the cone axis. A bluntness ellipticity (major-to-minoraxis ratio) of 2.0 was chosen because it has a

rather uniform heat flux distribution and relatively low stagnation point

heat flux (ref. 7). r

AERODYNAMICS

The Newtonian lift-to-drag (L/D) ratio is determined by the rake angle with

L/D approximately equal to cot ,i(ref. 8). In the present study the desired L/D of 0.3 requires a rake angle of 730 and appears to be sufficient

for control of the vehicle during atmospheric flight (r_fs. 9 and 10).

This configuration trims at an angle of attack of zero with respect to the

cone axis, thus placing the stagnation point on the maximum radius of

curvature of the ellipsoid.

Newtonian aerodynamic characteristicsof the concept are shown in figure 5.

At an angle of attack a = 0 the lift and drag coefficients are 0.45 and

1.53 respectively, resulting in an L/D of 0.296. The change in pitching

moment about the midpoint of the rake plane is Cm_- 0.0031 per degree.

The change in yawing moment _bout the same point due to sideslip angle is

Cn B- 0.0033 per degree. These coefficients represent a statically stable

vehicle.

.'EHICLECONCEPT

The layout of the vehicle was arrived at by balancing various requirements

and considerations. One ground rule for the study was that the vehicle be

transportable in sections in the Orbiter payload bay. This established the

maximum propellant tank diameter. A useable propellant capacity of 38 000

kg (84 000 ]b) was assumed which defined the total propellant tank volume.

For a given volume, a mlnimum tank fnassis achieved by using fewer, but

larger tanks. However, to reduce boiloff during an extended dwel] time at

7
GEO, one full liguid hydrogen fuel tank is dedicated to the descent propulsive maneuvers. This avoius having large partially filled tanks at GEO

in which the surface ared to volume ratio is not minimized. Heat transfer

through the large surface area would lead to excessive loss of hydrogen due

to boiloff. Boiloff is very costly due to the high price of transportation

to LEO from Earth. This arrangement is also convenient from a packaging

standpoint (fig. 6). Since boiloff is not as great a problem for liquid

oxygen, only one liquid oxygen tank is included in the design, and it is

used for both ascent and descent propulsion.

Two upgraded RL-IO engines having a t_rust of approximately 67 000 N

(15 000 Ib) each are included in the design. Two engines allow for a

failure oF a single engine without losing the vehicle.

The payload shown in fig. 6 as a cylinder is located as it would be for descent. It is thus pro ected from heating during aerobraking. A payload

that is to be delivered only to GEO may be attached to the vehicle at the

opposite side of the aerobrake from the engines, thus allowing for additional flexibility in payload configuration. In the present study the payload is assumed to be a 6800 kg (15 000 lb) manned capsule that is brought

back to LEO, although it is seen that a greater payload return capability

is possible from performance and aeroheating standpoints.

P_OPULSION

A propulsion system on an OTV utilizing an aerobrake must be integrated

with two other major elements, the aerobrake and the payload, rather than

just the payload as on an OTV without an aerobrake. Three different engine

firing arrangements are possible: side-firing, firing through the aerobrake, and firing away from the aerobrake. The last firing arrangement

would require the engines and propellant tanks to be placed around the periphery of a centrally located payload. If the payload is fairly long, exhaust from the engines will impinge upon the payload. Firing through the

aerobrake, like a conventional stage, would place the payload atop the propulsion system with the aerobrake addP_ below extendable nozzle engines.

The required doors in the aerobrake would add complexity and be unsafe for

manned missions, allowing a fatal single point failure. Therefore, side

firing would be the preferred arrangement for an aerobraked OTV.

To limit the size of the aerobrake, the propulsion system must be sandwiched between the payload on one side and the aerobrake on the other.

With different payload masses and propellant quantities, the center of

gravity of the vehicle will shift laterally with respect to the engine firing direction. A conventional stage would see little of this lateral shift

in the center of gravity. Therefore, greater engine gimbaling capability

will be required to maintain the thrust through the center of gravity.

Placement of the engines close to the line of center of gravity movement

and placing the oxygen tank and payload as far away from the engines as

possible (e.g. opposite side of aerobrake) may be required to minimize the

gimbaling. Also, the distance between the payload and the aerobrake should

only accommodate one tank, even though stacking the tanks will result in a

smaller aerobraSe. In additicn, the variable thrust direction may result

8
in an increase in propellant residuals because the propellant settles at

different angles with respect to the tank outlet. If future analysis determines that residuals need to be reduced, conical tank sections at the

outlet or acquisition screens may be required.

The placement of the propellant tanks on the aerobrake surface will determine the size of the aerobrake. A conventional tandem arrangement of tanks

with a single tank for each propellant would provide the minimum tank

weight and propellant boiloff while covering the minimum aerobrake surface.

However, the total area required for this arrangement with an aerobraked

vehicle having an elliptical plan form would be greater than evenly

distributing the tanks on an aerobraked vehicle having a circular plan

form. It was primarily this packaging problem that drove the design for

the aerobrake configuration to a raked elliptical cone, because the

projection of the shape into the rake plane was a circle whereas the

similar projection of a circular cone is an ellipse. Thus, the selected

configuration appears circular when viewed from the top, and this

configuration offered better packaging for the propulsion tankage. A

single cylindrical oxygen _ank and two cylindrical hydrogen tanks allowed a

conlpact,even distribution of tanks. The addition of a spherical hydrogen

tank filled the gap between the main tankage and the engines whileallowing

the boiloff to be minimized for the hydrogen required for descent from GEO.

With the spherical descent hydrogen tank, the boiloff would be minimized

because the surface area is at an absolute minimum. The added surface area

in the two cylindrical hydrogen tanks does not greatly affect boiloff

because of the limited time required to perform the ascent portion of the

mission.

Two RL-IO lIB engines were chosen for this concept. This upgraded version

of the engines used on the Centaur has a specific impulse typical of the

more advanced expander cycle engines; and the extendable nozzle is required

to prevent impingement of the exhaust on the aerobrake and also allows the

nozzle to be out of the high enthalpy entry flow. A minimum of two engines

are required for safety during manned missions, so that one engine can

always backup a failed engine.

VEHICLE STRUCTURE CONCEPT

To obtain the objective of a minimum mass structure for the total system

the integrated design was chosen as shown in figure 6. The propellant

ta_ks are supported by trunnions connected to web bulkheads which also form

part of the main thrust structure. Transverse bulkheads provide rigidity

normal to the thrust direction. The aerobrake structure is a lightweight

closed isogrid with a skin.

The engines are tied to the thrust structure with struts and the payload is

attached to the structure such that the loads are distributed uniformly

through four hard points.

Several gauges of structural materials were used in the design analysis.

Materials of various dimensions were used in different locations according

to load requirements. A schematic of the structural design is given in

figures 7 and 8.

9
STRUCTURAL MASS ESTIMATE

A Nastran model was made of the 12-meter (40-ft) diameter, b]unted,

raked-off conical AOTV design in order to make a preliminary mass estimate.

The heat shield was modeled using shell elements representing two different

patterns of closed isogrids. The engine and propellant tank mounting

struts were modeled using three sizes of rod elements. The bulkheads were

modeled using simple plate elements. The propellant tanks were modeled as

concentrated masses and rigid bars. Although the actual mass is less, the

thermal protection system was modeled as being 6.3 kg/m2 (1.3 lb/ft2) of

nonstructural mass applied to the heat shield shell elements. All other

components of the vehicle were accounted for with concentrated masses.

Three load cases were examined. Cases I and 3 were aerobraking descents at

3 g with no payload and with 6800 kg (15 000 Ib) of payload, respectively.

Load case 2 was a propulsive ascent at 2 g with a 28 DO0 lb payload and a

full propellant loaa of 39 000 kg (86 000 Ib). On an actual vehicle, powered flight at 2 g acceleration would occur only when the propellants were

nearly depleted. Therefore, this load case was extremely conservative.

A safety factor of 2 was used throughout the analysis. All elements had

positive margins of safety under all load cases. No buckling analysis was

performed and the bulkheads used on an actual vehicle would require stiffeners and might be replaced by open isogrid panels or truss assemblies.

This is not expected to change the mass estimate significantly.

Aluminum 7075 properties were used for all structural components. Tank

masses were estimated using aluminum 2219 properties. The structural mass

could be cut roughly in half with the use of composite materials, at the

price of increased manufacturing costs, additional design work, and possible service life limitations, depending on the material chosen.

The total estimated structure mass is 681 kg (1499 lb) composed of 375 kg

(825 Ib) for the heat shield structure, 206 kg (453 Ib) for the bulUheads,

and 100 kg (221 lb) for the struts. The total tank mass estimate is 1303

kg (2869 Ib) which includes thermal insulation, micrometeoroid protection,

baffles, and plumbing.

AEROBRAKING TRAJECTORY

Preliminary trajectory simulation using the 1962 standard atmosphere produced the trajectory plots shown in figures 9-12. Entry conditions were

10 311 m/sec (33 829 ft/s) velocity and -4.520 flight path angle to simulate return from GEO for an entry mass of 13 200 kg (29 000 Ib) and m/CDA =

73.8 kg/m2 (15.11b/ft2). The guidance scheme (ref. 9) computes reference

dynamic pressure and altitude rate to target the vehicle to a 370 km (200

nmi) apogee after atmospheric exit. The minimum altitude reached was 76 km

(250 000 ft), producing maximum dynamic pressure and maximum g-load of 1500

_ • N/m2 (32 psf) and 2.2 g respectively. The desired inclination of 28.5o was

achieved with four bank reversals.

10

-1 l
Preliminary analysis of vehicle performance in response to entry, aerodynamic, and atmospheric uncertainties shows adequate controllability with a

5°s"2 roll acceleration and a 17.5°s"1 maximum roll rate (ref. 11).

AERODYNAMIC HEATING ESTIMATES

The heat flux estimate follows that presented in reference 1. Heat flux to

the aerobrake windward surface is estimated using the trajectory

correlation

q R_N = 7.3 Im_ _C--_0"467 ) (_I 0"242 (W/cm2) mI/2 (i)

where

q = heat flux

RN = nose radius at stagnation point

m = mass of vehicle at entry

CD = drag coefficient

A = reference area

The heat flux used to obtain equation 1 is based on the engineering

correlation for the heat flux to the stagnation point of a reference sphere

q_r_N : 18 300 p® 1/2 (V_ /104)3.05 (W/cm2) mI/2 (2)

where

p_ = free stream density

V = free stream velocity

Using this correlation (eq. I) results in a small difference from the value

calculated from (eq. 2) when the actual trajectory conditions are used.

Through geometrical considerations the heat flux to the stagnation point

could be obtained. Using the above formulas with an aerobrake diameter of

12 m (40 ft) _nd mass of _3 200 kg (29 000 Ib) the peak heat flux was founa

to be 26 W/cm_ (23 Btu/ft -s) for a fully catalytic surface. Assuming a

factor C.75 to account for finite rate catalysis the estimated peak stagnation point heat flux is about 19 W/cm2 (17 Btu/ft2-s). This corresponds to

a temperature of 1430 K (2110°F) based on radiation equilibrium. This

temperature is well within the capability of the current Space Shuttle tile

material used on the lower surface of the Orbiter.

The heat flux to other areas of the aerobrake windward surface is assumed

to be no greater than the stagnation point value. Shock layer radiation is

11

.,%
assumed to be small (ref. 12) and is neglected. A representative heat flux

distribution for a similar body is shown in figure 13. This distribution

is calculated using a 3-D Navier-Stokes numerical solution (ref. 13). The

difference between the actual body and the one shown in figure 14 (used for

the calculations) is the handllng of the skirt. For the calcula+ion the

skirt is extended considerably and has a large radius of curvature. This

may have the effect of reducing the predicted skirt heat flux below a

realistic value.

THERMAL PROTECTION

The thermal protection system (TPS) envisioned for this study is an extension of Shuttle Orbiter technology. It is believed that there will be improvements in materials and concepts over the next few years, but these potential improvementswill be ignored in the present study. The capability

of the Shuttle high temperature reusable surface insulation (HRSI) tiles is

quite within the peak heating estimate of 19 W/cm2 (17 Btu/ft2-s). The

proposed TPS consists of HRSI tiles bonded with room temperature vulcanizing silicone rubber (RTV) to a strain isolation pad which is bonded to the

aluminum skin of the aerobrake. As in the Shuttle design, the tile thickness requirement is based on the aluminum skin not exceeding a temperature

)

of 450 K (350°F). This is a somewhat arbitrary requirement for an AOTV.

The tile thickness (approximately 12 mm) is based on the stagnation point

: heat load, and that thickness is assumed to be uniform over the entire

aerobrake. Thinner o{ lighter materials could be used in lower temperature

regions of the aerobrake and hence save weight; however, this refinement

was not incorporated here. Figure 16 of reference 7 is used in this study

to estimate the TPS mass per unit area, which was found to be

3.2 g/m2 (0.66 Ib/ft2) for an effective structure thickness of _ = 0.254

cm (0.1 in). For an aerobrake surface area of 146 m2 (1570 ft2), this

results in a total aerobrake TPS mass of 468 kg (1030 Ib).

TOTAL VEHICLE MASS

To obtain a total vehicle mass it was necessary to assume values for various subsystem components. These were estimated on the basis of current

state-of-the-art allocations and previous OTV studies.

A summary of the calculated and assumed masses is given in table 1. To

obtain the mass at reentry a factor of 1.0242 was applied to the end-ofmission estimated mass. This factor accounts for additional fuel on board

at the beginning of the atmospheric phase of the flight, subsequently used

I for circularization and LEO maneuvering.

i It can be seen from the table that the total aerobrake mass, including its

, thermal protection system, is about 17 percent of the total end-of-mission

stage mass excluding the 6800 kg (15 000 Ib) payload. This factor is much

smaller than previously estimated for aerobrakes that are attached to an

independent or nonintegrated stage and is approximately the same as that

estimated for a ballute concept (ref. 5). If one constructed the vehicle

12

)
from composite materials such as graphite polyimide or graphite epoxy, the

structural mass would be several hundred kilograms lighter and the aero-

_, brake mass fraction would be decreased to about 13 percent. Thus, the concept studied here is highly competitive based on mass considerations.

PERFORMANCE

Performance estimates for service to and from GEO were made based on the

following assumptions:

1. Impulsive burns

2. Isp = 4500 m/sec (460 ]bf-sec/lbm)with 1 percent degradation for

start and stop losses

3. Flight performance reserves assumed to be 2 percent of the & V

i" applied at each burn

4. _V assumptions:

For GEO missions

GEO transfer 2400 m/s (8000 ft/s)

Mid-course 15 m/s (50 ft/s)

GEO circularization 1800 m/s (6000 ft/s)

LEO transfer 1800 m/s (6000 ft/s)

Circularization at LEO

" after aerobraking 91 m/s (300 ft/s)

For lunar missions

Translunar 3150 m/s (10,350 ft/s) first stage

Mid-course 66 m/s (180 ft/s) second stage

Lunar circularization 970 m/s (3190 ft/s) second stage

Descent 2100 m/s (6890 ft/s) second stage

Ascent 1920 }n/s(6292 ft/s) second stage

Trans-Earth 947 m/s (3107 ft/s) second stage

Mid-course 55 m/s (180 ft/s) second stage

LEO circularization after

aerobraking 94 m/s (310 ft/s) first or second stage

4

g

5. Reaction control system propellants = 1 percent of main impulse

propellants

13
6. End-of-mission residuals = 438 kg (964 lb)

Using these assumptions the stage performance is

GEO delivery mass 13 500 kg (29 800 Ib)

Round-trip payload mass 7130 kg (15 700 Ib)

Mass ratio to GEO 4.334

Lunar surface delivery mass 18 800 k§ (41 500 Ib)

Round-trip to lunar surface 6800 kg (15 000 Ib) with 3860 kg (8500

]b) excess payload mass one-way to surface

Mass ratio to lunar surface 6.1

Figure 15 shows a comparison of this aerobraked stage with a similar all

propulsive stage with the same propellant load. The left hand ordinate

WGEO/WGEOAP is a measure of the additional payload delivered to GEO by an

aerobraked stage normalized by the GEO payload capability of the identical

all propulsive stage. The abcissa is the additional weight penalty of the

aerobraked stage. The plotted curve shows the percentage increase in GEO

delivery capability of an aerobraked stage as compared to an all propulsive

stage. Note that if the aerobraked stage was 50 percent heavier than the

equivalent propulsive stage, there is no performance improvement possible

with the aerobraked stage. However, as the weight penality for the aero

surfaces approach zero, the performance improvement approaches a maximum

limit of 96 percent. The design in this report, shown to be at the 17

> percent point, will yield a 76 percent increase in payload to GEO over the

all propulsive stage. Note that the approximation for an all composite

structure will yield an 80 percent improvement. Also note that this

design is satisfactorily close to the maximum yield of 96 percent . To

compare the aerobraked stage with alternative technologies, the right hand

ordinate is an indication of the Isp required of an all propulsive stage to

achieve the same improvement. An Isp of 5050 m/s (515 lbf-sec/Ibm) would

be required in order for an all propulsive stage to make the same GEO delivery; or, looking at the comparison in a different way, the aerobrake is

equivalent to a 540 m/s (55 lbf-sec/Ibm) increase in specific impulse. One

final comparison would be the costs saved by this aerobraked stage. This

comparison is made by using this aerobraked stage to deliver the 7600 kg

(16 800 ]b) all-propulsive-limit-payloadto GEO, then calculating the reduced launch costs for delivery of propellants to orbit. The aerobraked

stage can deliver the 7600 kg (16 800 lb) payload to GEO and save 9676 kg

(21 313 Ib) of propellants which equates to a $27.7 million savings in

Shuttle launch costs at $1,300 per pound.

7

14

A

• I
t

)

TECHNOLOGY IMPROVEMENTS

There are several areas of technology in which advances would lead to improved design and performance. Improved engines with greater specific impulse would reduce the total propellant and tank mass. Of course, the engines must be extremely reliable and this places a constraint on their design. Since propellant boiloff reduces that available for propulsion,

improved tanks could save propellant and reduce tank size. Reducing the

tank size would permit reducing the aerobrake size and thus reduce its

mass. The development of composite structural components would reduce the

mass of the structure without sacrificing strength or rigidity.

Since the thermal protection system considered in this study is essentially

minimum-gauge rigid insulation, the use of a lighter fabric or soft insulation having improved temperature and strength characteristics could reduce

the TPS mass.

Improved aerothermodynamic predictions of the heat flux distribution on the

aerobrake as well as in the wake or lee side regions of the vehicle will

enable tailoring the thermal protection system more accurately and thus reduce conservatism. Improved knowledge of the flow field chemistry and surface catalytic recombination reaction rates will improve heat flux predictions, as will improved numerical flow field computationalmethods. The

latter are particularly important for an AOTV since wind tunnels are limited in their simulat,on capability at these high altitudes and velocities.

CONCLUSIONS

This study of an integrated structural design of a lifting brake, aeroassisted orbital transfer vehicle has shown that it is possible to configure "i

a very mass-competitive space-based design. While this rigid design must

be assembled in orbit from subassemblies transported in the Space Shuttle

payload bay, it has the advantage of not requiring the replacement or

repackaging of major components such as an inflatable structure.

The ellipsoidally blunted, raked-off elliptical wide-angle cone has several

desirable features, but it should be compared with other lifting brake

shapes such as a blunted symmetrical cone with an offset center of gravity.

Wind tunnel and computational fluid dynamic analyses will help assess the

aerodynamics and aerothemodynamics of each configuratiun.

For the size of return payload and the tank sizes considered the thermal

requirements imposed on thermal protection system are fairly mild and can

be satisfied by current technology. This implies that with smaller tanks

and return payloads one could reduce the diameter of the aerobrake and significantly reduce its mass without exceeding TPS temperature limits due to

excessive aeroheating. This study has assumed that the nonequilibrium

shock layer radiative heat flux is very small and is consistent with the

calculations of reference 12. If the nonequilibrium radiation were considerably larger as implied by reference 7, then the TPS would be somewhat

heavier. With minor improvements in current TPS materials' capability

(e.g., a more reflective coating) significant radiative heating can be

accommodated.
TABLE I - JSC AOTV CONFIGURATION MASS SUMMARY (GEO RETURN)

Mass (in pounds)

TPS 1 030

Heat Shield Structure 825

Aerobrake subtotal 1 855 (16.7% of stage at EOM)

Bulkheads 453

Struts 221

Structure Subtotal 1 499

Stage - payload 585

Internal insulation 526

Power and distribution 480

RCS 449

Avionics 464

Tankage 2 869

Engines (two RL-IO) 1 855

Total dry stage mass

Residual & reserve

propellant (EOM) 1 337

End of Mission Mass (Stage) 11 094 Ib

Contingency 20% 2 219

Payload (manned capsule) 15 000

Total EOM mass

Reentry mass 1.024 x EOM mass 28 993 lb

m

_ = 15.1 lb/ft 2
100 -

O_

O0

f_

ALL PROPULSIVE

" O ;_B- 4

LU

" rO

_: .3

O

LU AEROL'RAKED

-J

_: .5

10-

¢

• . . . I 1_._ 1 I,., I I 1_ 1. I. I I

" .7 .8 .9

MASS FRACTION, X

Figure 1. - Payload-to-stage mass ratio of a11 propulsive and

aerobraked stages.
18

,,.
I

ELLIPSOIE

RAKE PLANE

/

/

/

Ftsure 3. - Geometric construction of blunted,

raked-off cone.
oRIGtNAL P':_'-

OF pOOR QUAL|'i_f

, VELOCITY

, •

A

Ftsure 4. - Integrsted AOTV design in aerobrsklu$ attitude.

2O
_.8 m

1.6

1.4

_ 1.2

m 1.0

E

o

"n

.2

0

-0.2

-20 -15 -10 -5 0 5 10 15 20

ANGLE OF ATTACK, DEG.

Figure 5. - Nevtonian aerodynamic coefficients of blunted,

raked-off cone.

21

i_ p
i

OF POO,_ QL,/,,Li'C_

-; TOP VIEW

ASCENT DESCENT

HYDROGEN HYDROGEN TANK

TANK

OXYGEN TANK j ENGINE

ASCENT

HYDROGEN

TANK

'ENGINE

b

HEAT SHIELD

WIND SIDE VIEW

VECTOR

GIMBAL ANGLE =

THRUST

VESTOR

,PAYLOAD

i Figure 6. - Integrated AOTV concept with 12-m (40-ft) diameter

heat ahleld.
X

23

i

40

/

_1' //

._ 34 /

ui 32

a

-J

< 30

28 I

/

24

0 50 100 150 200 250 300 350 400 450 500 550

TIME, SEC

Flgure 9. - Altitude tlme hlstory durlng AOTV aerobraklng.

25
'l

-t

: 34 _ ._

:' 33 3[_

32

31

i ¢n

_- 30

u.

• .lg

,_, _ 29

, o \ i 28

27 ,_

26 _ " "-'-'_ _ eO

25 I ..

0 50 1O0 150 200 250 300 350 400 450 500 550

TIME, SEC

Figure 10. - Velocity time history during AOTV aerobrakin 8,

26
2OO

: uJ 100

° J '

ui

_J

<Zv 0 "-'-:'a-_ ] f _=-:, -1O0 _- #_

z< .__,.w% ._jl -200 , , , 1 ,,,

0 50 100 150 200 250 300 350 400 450 500 550

TIME, SEC

(a) Bank angle

_. --G-E-E. ,--c--,-'(3E) _

(.9

° /

T

_" -2

-J -4 ....

LL

-6 t,,, .... I 7 1 , ,

0 50 100 150 200 250 300 350 400 450 500 550

TIME, SEC

(b) Flight path angle

29j ¢N

JJ

° j/

o 28

g

z_

_J

27 _'

Z

26 '

0 50 100 150 200 250 _100 350 400 450 500 550

TIME, SEC

(c) Orbital inclination

Figure ii. - AOTV atmospheric flight trajectory

characteristics.

27
40 -,

I--

I1

.tn 30 -- // _ _

Lu"

fr

2O

OO

Og

LU

° I....1 J I

z

>- 0 50 100 150 200 250 300 350 400 450 500 550

{3 TIME, SEC

(a) Dynamic pressure

- 3

o_o S_, \

.J

°' 2 "-" Ii

0 50 100 150 200 250 300 350 400 450 500 550

TIME, SEC

(b) Atmospheric deceleration

,,xJ:D 150 /_

o--U'II--_UJ 100 _I'I" /-- ._X l

0 50 100 150 200 250 300 350 400 450 500 550

TIME, SEC ,

(c) Heat flux to reference 1-FT sphere

. F$sure 12. -AOTV atmospheric flisht environment.

w

4

28

4
V== = 9 75 km/s M= = 34 8 e = 0 IDEAL GAS p_= = 7 47 x 10-6 kg/m 3 & = 75°

(32 kft/s) exy = 60° _b = 2 RN = 1 52 m (1 45 slug/ft 3)

(5 ft)

1 0 _- 1.0 -

8 .8

_o o

< .6 .6

oE fie

x

.,../

< _ .4

.2

"" 0 ' I I I 1 0 I I I I

-10 -5 0 5 10 -10 -5 0 5 10

Y-AXIS, FT Y-AXIS, FT

Figure 13. - Heat flux and pressure dlstrlbuClons on aerobrake

referenced to scagnatlon point value.

29

q ,t, p,
Y

\

Figure 14. - Computational body with Math number

contours in shock layer.

3O
Ibf - sec m

Ib m sec

1.0

5200

COMPOSITE STRUCTURE

.8 -80% (67% CONVENTIONAL)

-76% DNAL STRUCTURE 5100

¢/)

WGEOA P -- 16 800 LB 5000 w ,6 - >_

WSA P = 9 244 LB 4900 (JW

n U.

,o,_ Ap --0.9Ol "W

0 _ .4 - 4800 uJ

<! C9

<C

I--

4700 u')

.2 - a

w

4800 _'

<[

m

O

0 4500 n."

uJ

,<

12.7% 17%

0 .1 .2 .3 .4 .5 .6

kAB, AEROBRAKE WEIGHT PENALTY

Fisure 15.- Fractlonal increase of GEO delivery mass for

aerobraked stage over an all propulsive stage.

31
REFERENCES

1. Scott, C. D.; Ried, R. C.; Maraia, R. J.; Li, Chien P.; and Derry, S.

M.: An AOTV Aeroheating and Thermal Protection Study. AIAA Paper

84-1710, June 1984.

2. Hethcoate, J. P.; and White, J.: Aeromaneuvering Orbit Transfer Vehicle

for Space Transport System, Paper IAF-76-173 presented dt XXVII

Congress InternationalAstronautical Federation (Anaheim, Calif.),

Oct. 10-16, 1976.

3. Walberg, G. D.: A Review of Aeroassisted Orbit Transfer. AIAA Paper

82-1378, Aug. 1982.

4. Walberg, G. D.: Aeroastisted Orbit Transter-WindowOpens on Missions.

Astronautics & Aeronautics, vol. 12, no. 11, Nov. 1983, pp. 36-43.

5. Andrews, D. G.; Caluori, V.A.; and Bloetscher, F.: Optimzation of

! Aerobraked Orbital Transfer Vehicle. AIAA Paper 81-1126. Also

I published in Thermodynamics of Atmospheric Entry, AIAA Progress in

,| Astronautics & Aeronautics, vo]. 82, 1982, pp. 455-476.

: 6. Florence, D. E.: AerothermodynamicDesign Feasibility of a Generic

i Planetary Aerocapture Aeromaneuver Vehicle. AIAA Paper 81-1127.

i Also published in Thermodynamics of Atmospheric Entry, AIAA Progress

' in Astronautics and Aeronautics vol. 82, 1982, pp. 477-519.

i

4 7. Menees, G. P.; and Park, C.: Design and Performance Analysis of a

Conical-AerobrakeOrbital Transfer Vehicle Concept. AIAA Paper

84-0410, Jan. 1984.

!

i 8. Mayo, E. E.; Lamb, R. H.; Romere P. 0.: Newtonian Aerodynamics for

, Blunted Raked-Off Circular Cones and Raked-Off Elliptical Cones, NASA

TN D-262, May 1965.

9. Gamble, J.; Cerimele, C.; Spratlin, K.: Aerobraking of a Low L/D Manned

Vehicle from GEO Return to Rendezvous with the Space Shuttle, AIAA

Paper 82-2110 presented at AIAA Atmospheric Fli]ht Mechanics

Conference (Gatlinburg,Tenn.), Aug. 15-17, 1983.

10. Cerimele, C.; Skalecki, L.; Gamble, J.: Meteorological Accuracy

Requirements for Aerobraking Orbital Transfer Vehicles, AIAA Paper

84-0030 presented at AIAA 22nd Aerospace Sciences Meeting (Reno,

Nev.), Jan. 9-12, 1984.

11. Gamble, J.; Spratlin, K.; Skalecki, L.: Lateral Directional

Requirements for a Low L/D Aeromaneu,,eringOrbital Transfer Vehicle.

AIAA Paper 84-2123 presented at AIAA Atmospheric Flight Mechanics

Conference (Seattle,Wash.), Aug. 1984.

32
12. Ried, R. C.; Rochelle, W. C.; and Milhoan, J. D.: Radiative Heating to

the Apollo Command ModuIe_ Engineering Prediction and Flight

Measurement, NASA TM X-58091, Apr. 1972.

13. Li, C. P.; A Three Dimensional Navier-Stokes/EulerCode for Blunt-Body

Flow Computations, AIAA Paper 85-0361 to be presented at the 23rd

AIAA Aerospace Sciences Meeting (Reno, Nev.) Jan. 1985.

33