NASA/TM--2002-211599

SRM-Assisted Trajectory

GTX Reference Vehicle

for the

John Riehl and Charles Trefny

Glenn Research Center, Cleveland, Ohio

Daniel Kosareo

ZIN Technologies, Inc., Brook Park, Ohio

June 2002
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NASA/TM--2002-211599

SRM-Assisted Trajectory

GTX Reference Vehicle

for the

John Riehl and Charles Trefny

Glenn Research Center, Cleveland, Ohio

Daniel Kosareo

ZIN Technologies, Inc., Brook Park, Ohio

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

National Aeronautics and

Space Administration

Glenn Research Center

June 2002
The Aerospace Propulsion and Power Program at

NASA Glenn Research Center sponsored this work.

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Hanover, MD 21076

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

ZlN 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 Ib 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 1
SYMBOLS

A*

Ac

Isp

M

Po

P4

Pc

q

T

Wf

Y

Rocket Engine Throat Area

Vehicle Inlet Capture Area

Specific Impulse

Mach number

Atmospheric Pressure

Duct Pressure

Rocket Engine Chamber Pressure

Dynamic Pressure

Net Thrust

Final Weight

Angle of attack (deg)

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

ACRONYMS

OTIS

RJ

RAM

SRM

SC

SCRAM

Optimal Trajectories by Implicit Simulation Computer Program

Ramjet Engine

Ramjet Engine

Solid Rocket Motor system

Supersonic Combustion Ramjet Engine

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 (R J)

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 2
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, (z, 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, (z, 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, (z 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 3
definition for the optimizer. Lastly, the initial azimuth after pitch-over, _, is also open for

optimization, but with 0 < • < 90°.

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 Ib/ft 2. The minimum dynamic pressure for

RJ/SJ mode is 500 Ib/ft 2 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 Isn is 277.8 secs. and the

total exit area is 22.3 ft 2. 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*PcfJM, P--_"

t Po

l"t '7oJ

NASA/TM--2002-211599 4
where A* is thruster throat area in ft2, Pc is chamber pressure in Ib/in 2,Po is atmospheric pressure,

and the functionsfil andS1 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, A, is

138.2 ft2 and A * is .942 ft2. c

Modes 2 and 3 (Ramjet and Scramjet)

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

T=qAJvN(M)

I_P= f,N(M)

P4 = f,N(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 aSfix and_x

where N=2 or 3.

Mode 4 (Rocket)

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

chamber pressure.

L ....... =A*Pfr4(P)

isp:

Here againfi4 and_ 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 = T_....... -- A_i_Po

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 5
Gravitational potential was assumed to be 1.40785E+16 ft3/sec 2 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 6
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 Ibs. 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 7
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 Maximum Minimum Duct Maximum Duct

Mach Mach Pressure (Atm.) 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

Freestream dynamic pressure, Qo (psfa)

Angle-of-attack, (z (degrees)

Thruster chamber pressure, Pc (psia)

Diffuser exit pressure, P4 (atm)

Total acceleration (g's)

Mode 1

Min Max

500 1500

400 2000

4

Mode 2

Min Max

500 1500

-6 6

1/2 10

4

Mode 3

Min Max

500 1500

-6 6

1/2 10

4

Mode 4

Min Max

500 1500

400 2000

NASA/TM--2002-211599 8
Table 3. KEY MISSION EVENTS FOR THE BASELINE AND THRUST AUGMENTED VEHICLES

Flight

Phase

Liftoff

Time

(sec)

0

0

Altitude

(ft) (Nmi)

0

0

Relative

Velocity

(ft/sec)

0

0

Mach Weight

(Ib)

236,000

360,770

End of 6.1 500 0.08 163 0.15 229,027

Vertical 5.35 500 0.08 191.4 0.17 347,003

Rise

Mode 1-2 65.1 43,015 7.1 2414.5 2.5 174,017

53.5 43,015.3 7.1 2414.5 2.5 240,633

End SRM N.A. N.A. N.A. N.A. N.A. N.A.

80 62,696.2 10.3 3801.5 3.9 217,134

then

203,700

Mode 2-3 149.6 78,277 12.9 5,374 5.5 167,455

114.3 72,242 11.9 4,868.6 5 201011

Mode 3-4 613.5 122,443 20.2 11,183 10.9 133,314

617.3 112,838 18.6 10,821 10.7 169,659

End of 1st 720.3 228,321.5 37.6 24,568.5 25.1 50,996.5

rocket burn 731.8 217,737.4 35.8 24,673.2 24.7 59,937.5

Start 2nd 2,201.8 1,340,829.8 220.7 23,160.4 20.7 50,996.5

rocket burn 2,421.8 1,322,160.8 221.2 23,245.3 20.8 59,937.5

End 2,209.8 1,340,940.2 220.7 23,731.8 21.3 48,827.9

2,429.9 1,344,223.2 221.2 23,729.4 21.3 57,769.1

Table 4. PROPELLANT BREAKDOWN FOR THE BASELINE AND

THRUST AUGMENTED VEHICLES

Engine Mode Oxidizer Weight Fuel Weight (lb.) Total Propellant O/F Ratio

(lb.) (lb.)

Mode 1 52,038 9,533 61,571 6.34 (aver.)

24,693 5,415 30,108 5.90 (aver.)

Modes 2 and 3 0 34,125 34,125 N.A.

0 36,104 36,104 N.A.

Mode 4 79,003 11,671 90,674 6.77

97,207 14,360 111,567 6.77

Total 131,041 55,329 186,370 2.37

121,900 55,879 177,779 2.18

NASA/TM--2002-211599 9
Figure 1. Isometric view of GTX with SRM added

.i_i__,i'°° ,_C_E_3o_.

li1 _ io_

1

SRM-_.SSfSTE_GT_ iOC

Figure 2. Three-View of GTX 10C with GEM 46SRM's

NASA/TM 2002-211599 10
GEM 46 Thrust Profile

(thfe_ en_inesl

S"

400000

E

300000

200000

700000 :-.......................................................................................................................

OTIS mo:del

tn#u_:to Model

100000

40 20 30 40 50 60 70 80

Time {sees)

Figure 3. GEM 46 thrust profile

G

10

8

6

4

2

0

_2

_4

_6

_8

Thrust Aug_mented

............E_aselir_e

0 5 10 t5 20 25 30

Math

Figure 4. Angle of Attack vs. Mach

NASA/TM 2002-211599 11
90.

.I _ Thru_I A_Jgmer_ed

I ....................... gaseline

_0 " ............................................................................................................................

7O

6O

e

._ so

N

._ 40

¢-

30

20

I0

0 5 10 15 20 25 30

Mach

Figure 5. Flight Path Angle vs. Mach

350000

30000:0

-- Thrust: Augmented

.......................... Base_ir_e

25000:0

g 200000

150000

100000

50000

0 500 1000 1500 2000

Tim_ (sees}

Figure 6. Weight vs. Flight Time

2500

NASA/TM 2002-211599 12
250

200

Is0

g

< IO0

50

..................................................... ;,;j,_'-_2 _-.........................

,S" ,,,."

.<<' _,, .............. Base F_e

/o/

......... iii .............................................................................

i I I

0 2000 4000 6000 8000

Dow_ Range Oistm_ce {Nmi)

Figure 7. Altitude vs. Down Range Distance

250

200

I50

lOO

50

5 10 15 20 25 30

Mach

Figure 8. Altitude vs. Mach

NASA/TM 2002-211599 13
2500

E

0

2000 _""_

i000 ............................................

"]"hr[_St A_ugm_nted

......................... 13a,,3elil_e

500 I----lllI_-_-_ .................................................................................... _.......................

0 _ _............. ,_4,_,............................ _,,. _ _ _-,,,,_,t_, ............... _ ..............

0 5 i0 15 20 25 30

Math

Figure 9. Chamber Pressure vs. Mach

4,,5

@

<

4

3,5;

3

1.5

0,5;

Thrus_: Au#mented

......................... Baselir_e

iii i;iiiiii;i_ ................. i;.........

0 5 I0 15 20

Mach

Figure 10. Acceleration vs. Mach

NASA/TM 2002-211599 14
1400..... _.....

Thrust Augmented

............. Baseline

0.

U

o

1200

1000

800

600

400

200

,\

:<

0 5 I0 15 20 25

Math

Figure 11. Dynamic Pressure vs. Mach

30

900000

_ Thrust Aug:mented

.......................... Baseline

800000

70000:0

600000

500000

2

400000

300000

200000

1Q 000:0

5 10 15 20

Mach

Figure 12. Total Vehicle Thrust vs. Mach

25 30

NASA/TM 2002-211599 15
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Figure 13. Effective Isp vs. Mach

NASA/TM 2002-211599 16
Form Approved REPORT DOCUMENTATION PAGE

OMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

June 2002 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

SRM-Assisted Trajectory for the GTX Reference Vehicle

6. AUTHOR(S)

John Riehl, Charles Trefny, and Daniel Kosareo

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU-708-90-63-00

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-13386

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--2002-211599

11. SUPPLEMENTARY NOTES

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.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Categories: 15 and 07 Distribution: Nonstandard

Available electronically at http:/Igltrs.grc.na_sa.gov/G[2I'RS

This publication is available from the NASA Center for AeroSpace Information, 301-621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

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.

14. SUBJECT TERMS

Airbreathing propulsion; Launch vehicles; Hypersonic propulsion; Trajectory analysis

and optimization

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

NSN 7540-01-280-5500

15. NUMBER OF PAGES

22

16. PRICE CODE

18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT

OF THIS PAGE OF ABSTRACT

Unclassified Unclassified

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39-18

298-102