NASA/TM-2001-209620

A Two-Stage-to-Orbit Spaceplane Concept with

Growth Potential

Unmeel B. Mehta and Jeffrey V. Bowles

February2001

https://ntrs.nasa.gov/search.jsp?R=20010041080 2018-05-15T07:01:33+00:00Z
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NASA/TM-2001-209620

A Two-Stage-to-Orbit Spaceplane Concept with

Growth Potential

Unmeel B. Mehta

Ames Research Center, Moffett Field, California

Jeffrey V. Bowles

Ames Research Center, Moffett Field, California

National Aeronautics and

Space Administration

Ames Research Center

Moffett Field, California 94035-1000

I

February2001
Acknowledgments

While at the NASP National Program Office, Unmeel B. Mehta advocated that the NASP program should set aside

the SSTO requirement and develop a TSTO spaceplane with air-breathing propulsion on the first stage. Subsequently, Thomas J. Gregory directed at NASA Ames Research Center the development of the TSTO spaceplane

concept for the NASA access to space study. The authors thank Thomas J. Gregory (retired), Howard E. Goldstein

(retired), William D. Henline, and David J. Kinney of NASA Ames Research Center, William J. D. Escher of

Scientific Applications International Corporation, and Henry G. Adelman of ELORET Corporation for their helpful

comments and suggestions when they reviewed the Manuscript of this paper.

This paper is presented as AIAA Paper 2001-1795 at the AIAA/NAL-NASDA-ISAS 10th International Space

Planes and Hypersonic Systems and Technologies Conference, 24-27 April 2001, Kyoto, Japan.

NASA Center for AeroSpace Information

7121 Standard Drive

Hanover, MD 21076-1320

(301) 621-0390

Available from:

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22161

(703) 487-4650
A TWO-STAGE-TO-ORBIT SPACEPLANE

CONCEPT WITH GROWTH POTENTIAL

Unmeel B. Mehta and Jeffrey V. Bowles

NASA Ames Research Center, Moffett Field, California, U.S.A.

Abstract

A two-stage-to-orbit (TSTO) spaceplane concept

developed in 1993 is revisited, and new information is

provided to assist in the development of the nextgeneration space transportation vehicles. The design

philosophy, TSTO spaceplane concept, and the design

method are briefly described. A trade study between

cold and hot structures leads to the choice of cold

structures with external thermal protection systems. The

optimal Mach number for staging the second stage of

the TSTO spaceplane (with air-breathing propulsion on

the first stage) is 10, based on life-cycle cost analysis.

The performance and specification of a prototype/

experimental (P/X) TSTO spaceplane with a

turbo/ram/scramjet propulsion system and built-in

growth potential are presented and discussed. The

internal rate of return on investment is the highest for

the proposed TSTO spaceplane, vis-h-vis a singlestage-to-orbit (SSTO) rocket vehicle and a TSTO

spaceplane without built-in growth. Additional growth

potentials for the proposed spaceplane are suggested.

This spaceplane can substantially decrease access-tospace cost and risk, and increase safety and reliability in

the near term. It can be a serious candidate for the nextgeneration space transportation system.

Introduction

In 1959, the request for recoverable booster system,

with a goal of routine access to space, led to the

Recoverable Orbital Launch System (Aerospace Plane)

program in the United States (U.S.). _The U.S. Air

Force emphasized two-stage-to-orbit (TSTO) concepts

as first-generation options, based on guidance from the

Air Force Scientific Advisory Board (SAB) and other

ad hoc committees. In November 1965, after intensive

study, review, and evaluation, the TSTO approach (airbreathing first stage with a conventional rocket second

stage) was selected as the preferred approach. In 1970,

when the U.S. Space Shuttle Phase B award began, the

National Aeronautics and Space Administration

(NASA) and contractors were generally unanimous in

considering TSTO fully reusable vehicles as the vehicle

of choice. 2 In 1993, RAND believed that the 1965

choice by the U.S. Air Force was commendable and

that it would be a strong contender for developing the

X-30 spaceplane under the National Aero-Space Plane

(NASP) program, t In 1993, after reviewing the status of

X-30, the U.S. General Accounting Office (GAO)

recommended a re-examination of the worth of pursing

single-stage-to-orbit (SSTO) on its own merit. _ In 1994,

Ben Rich expressed that Orient Express (X-30) is

actually two separate concepts, one a rocketship and the

other an airplane. _ "Most likely, that particular twain

shall never meet successfully. ''a Based on state-of-theart hypersonic technologies) '__ the SSTO concept is an

extreme technical challenge, and its commercial

viability is highly questionable in the near term.

Nevertheless, the SSTO concept is still advocated. 7

Today, the choice of TSTO concept as the near-term

option, with initial operational capability (IOC) in circa

2012, appears to be the correct one. In 1993, a

spaceplane based on the TSTO concept was

conceptualized for the Access to Space--Advanced

Technology Team. 8Reference 9 briefly describes this

spaceplane. Reference 10 presents a strategy for

developing air-breathing spaceplanes, using a TSTO

spaceplane concept. Essential aspects of this concept

and of this strategy are revisited, and additional

information, some based on subsequent work, is

provided herein for consideration to assist in the

development of next-generation spaceplanes.

The lessons drawn from past programs suggest the

following design philosophy for the development of the

next generation space transportation system: 1_(I)

technology-driven development, (2) short-term

economical benefits, (3) growth potential, (4)
achievable, and(5)safeandreliable,evenatthe

expenseofgreaterup-frontcostandlowerpertormance.

Theeconomical benefitsarejudgedfor

commercialization ofspacetransportation. By

definition, acommercial ventureisalow-risk

investment andgeneratesa20-plus-percent returnper

annumoninvestment withtheinvestment payback

(preferably) within5years.Thedesignphilosophyand

economicsrcquirement leadtoaspaceplane concept

basedontwostages, toaspaceplanethatoperateslike

anaircraft,andtoaspaceplanc designwithbuilt-in

growthpotential.

Stagingcanincreaseperformanceofagiven

technology, rcduccthevehiclesensitivityto

performanceparameter variations, ordeliverequal

pertbrmanceandloweriskwithlessadvanced

technology. TSTOvehiclesoffergreatermarginand

havehigherpayloadpotential thanSSTOvehicles.

TSTOvehiclesrequiresmallerpropellant mass

fractions, resultinginlowertake-offgrossweight

(TOGW)thanSSTOvehicles. If vehicleswithairbreathingpropulsionareconsidered, theusefulairbreathingcorridorforTSTOvehiclesislargerthanthat

forSSTOvehicles, andfirststagesofaTSTOconcept

havepotential forgreateratmospheric-cruise capability

thanSSTOvehicles.Theneedfordevelopingtwo

stagesiofasecondaryconsiderationtothe

aforementioned designphilosophy, economic

requirement, andadvantages overSSTOvehicles.

Orbital-missionflexibilityandgreatlyenhanced

operabilityareachievable, if spaceplanes havefeatures

thatapproachthoseofcommercial aircraft.Airbreathingpropulsionprovideshigheroverall

performanceandfargreateroperabilityhanthat

possiblewithrocketpropulsion. Examplesofenhanced

operabilityareaircraft-likepowered-landing operations,

includingo-aroundandvehicleintact-abort capability.

Inadditiontotheaforementioned attributes, thefull

reusabilityof spaceplanes leadstosignificantlyreduced

operational costs,whichinturn,reducethelife-cycle

costs(asumofdevelopment, acquisition, andoperation

costs)ofafleetofspaceplanes. Ifpaceplanes are

designedwithbuilt-ingrowth,life-cyclecostsare

furthereduced. TheTSTOspaceplane, withairbreathingpropulsiononthefirststage,canfulfillthe

near-termaccess-to-space goals,whilegreatlyreducing

thecost,substantiallyimprovingmissionflexibility,

operability, safety,andreliability,andoffering

significant growthpotentialsandmultipleavenues.

TSTO Spaceplane Concept

Figure I shows an artist's rendition of the TSTO

spaceplane considered for the mission of achieving 220

nautical miles circular orbit at a 51.6 ° inclination, with

25,000-1b payload [International Space Station Alpha

(ISSA) resupply mission]. The genesis of this

spaceplane is the Study of Advanced Air-breathing

Launch Vehicles with Cruise Capability by Lockheed

Company in 1967 for NASA Ames Research Center) _

Figure 1. Artist's rendition of the TSTO

spaceplane concept. 8
Both the first stage (launcher) and second stage

(orbiter) vehicles are lifting-body configurations. The

orbiter is nestled within the outer mold line of the

launcher. The cargo bay dimensions of the orbiter are

15 x 15 x 30 ft. The launcher propulsion system

consists of a turbo/ram/scramjet propulsion system,

including low- and high-speed air-breathing engines,

while the orbiter propulsion system is a conventional

liquid oxygen (LOX)/hydrogen rocket engine.

The structural concept for the launcher and orbiter is

cold-structure, skin-stringer/frame-stiffened, integrallobed tanks (aluminum-lithium alloy for LOX tank and

composite for hydrogen tank), with composite intertank structure. The external thermal protection system

(TPS) consists of advanced ceramic tile/blanket

systems. Both stages of the spaceplane are designed

with 15-percent dry-weight margin.

Each stage of the spaceplane has a crew of two. The

launcher takes off horizontally and both stages land

horizontally, with the launcher making a powered

landing. The take-off speed is limited to 300 knots.

Staging is executed at low hypersonic Mach number

(5 < M < 12). The maximum axial acceleration during

ascent is limited to 3.0g, and the normal load factor is

limited to 2.5g.

After launching the orbiter, the launcher executes a

decelerating 180-degree turn and cruises on the lowspeed air-breathing system back to the launch site.

After separation, the orbiter continues to accelerate up

to main engine cut-off (MECO), coasts to the target

apogee, and circularizes at that altitude. The orbiter has

an on-orbit AV of 600 ft/s. After completing the onorbit mission, the orbiter de-orbits, enters the

atmosphere, decelerates aerodynamically while banking

to meet cross-range requirements, and executes an

unpowered landing at the launch base.

A study was conducted to determine the extent to which

the orbiter is submerged within the launcher. Structural

depth of the launcher in the payload region is traded for

transonic drag increment of the mated combination. The

lowest weight system resulted when the orbiter was

fully submerged.

Hydrogen is the preferred fuel for the launcher for a

number of reasons. First, hydrogen offers significant

propulsion efficiency and thermal cooling capability

over hydrocarbons, resulting in lower weights and

costs. Second, the sizing of the hydrocarbon-powered

launcher is principally determined by the size of the

orbiter, whereas the size of the hydrogen-powered

launcher is mainly determined by its fuel-fraction

mission requirements. In the latter case, the possibility

of forming hydrogen on-aboard from hydrocarbons and

water is not considered. If it were considered, the onboard hydrogen generation Iz would also impact the

sizing of the launcher. Third, hydrocarbons limit the

launcher performance growth. Fourth, the staging Mach

number can be higher with hydrogen than with

hydrocarbons. Finally, there is little risk-avoidance

provided by hydrocarbons on the launcher if hydrogen

is used on the orbiter.

Hydrogen-powered, after-burning turbojets were

selected for the launcher low-speed propulsion system.

Figure 2 conceptually shows the jet engines installed on

the launcher, consisting of an over-and-under

arrangement of turbojets and ram/scramjets. As a

function of flight Mach number, the inlet splitter plate

position is set to provide the turbojet engine the

required inlet mass flow. The turbojet engine sea-level

thrust level is in the 70,000-pound class. Six turbojet

engines are installed; they are used from Mach 0 to 2.4,

including takeoff and transonic acceleration. The

weight of the turbojet system includes bare engine

weight, afterburner and nozzle weight, and the close-off

door, resulting in an installed engine thrust-to-weight

ratio of 6.65. The operating range of turbojet engines

could be extended to Mach 3.5, with present, state-ofthe-art turbojet technology.

__ Pivoting flow _.. Turbojet,high-speed research

splitter _ propulsion(HSRP)technology

_ _ Pivot_ _ levelor F-22 engine

 0.w 'z.:21, I, ,1970 technology,

1/10 heat flux Mechanicalthroat

Figure 2. Jet engines installation. 9

The ramjet is started at Mach 1.05, to cancel ram drag

and internal friction, and operated to 6.0. From Mach

0.9 to 1.4, base burning is used to reduce transonic

nozzle drag. At the transonic pinch point, the resultant

acceleration is 2.3 ft/s 2. The reduced transonic specific

impulse (I_p), shown in figure 3, reflects the fuel

consumed for base burning. Above Mach 6, the NASPderived, dual-mode, ram/scramjet engine technology

is used.
Theorbiterpropulsionsystemconsistsof main rocket

engines and an integrated OMS/RCS system. The

baseline OMS/RCS system is a liquid H2/O2-fueled

system. The main rocket engines are expander cycle,

H2/LOX engines, based on RL-10 technology. The

thrust class of the engine is approximately 150,000 lb,

with an engine thrust-to-weight ratio of 54. It was well

within 1993 technology limits and is similar to the Pratt

and Whitney RL2000-study engine. Four engines are

selected to provide abort-to-orbit capability from the

launch point, if an engine is shut down. From Mach

5-launch condition, the orbiter can attain an orbit

altitude of 65 nautical miles, which is sufficient to

return to the launch site on the next orbital pass.

1.0

o

t_

n

0.5

W

0 1 2 3 4 5

Mach Number

Figure 3. Launcher propulsion-specific impulse.

A low-pressure, expansion-cycle rocket engine is used

due to its simple architecture, which would result in

high operational reliability and maintainability, and in

relatively low development and acquisition costs. In

fact, the orbiter could use any of the advanced rocket

engines, such as an advanced-staged, combustion-cycle

engine, modular engines, aerospike engines, etc. The

payload capabilities of the orbiter could be enhanced

with these advanced, higher-performance, lower-weight

engine systems.

Analysis Method

The ARC-developed hypersonic aerospace vehicle

optimization code (HAVOC) for hypersonic vehicle

synthesis is used to perform estimates of systems

performance characteristics and design specifications

(fig. 4). It uses engineering analysis methods to

compute vehicle performance and design

characteristics, including aero- and thermo-dynamics,

propulsion, structures, trajectory, and system cost. In

the design process, the synthesis code closes the design

to meet mission performance requirements, matching

vehicle weight and volume required to that available.

I COf_

[

Ico-.o,.-I

i

I oo.o.,o° I

TROL I

I I

I...s,.v,-I I i

I

II I I..oPo ,o.I IsT.ocTu. s

Figure 4. Hypersonic aerospace vehicle optimization

code (HA VOC).

Aero- and Thermo-Dynamics

The subsonic and transonic aerodynamics for a given

configuration are computed as functions of the free

stream Mach number, angle-of-attack, and the gross

geometric parameters of the body and wing (fig. 5).

Experimental or high-fidelity computational results are

used to calibrate the empirical methods in the lowspeed regime for the body, including the high-subsonic/

transonic/low-supersonic base drag model. In the

supersonic/hypersonic speed range, the fuselage

pressure and aerodynamic forces (lift and pressure

drag) are computed using real-gas tangent-wedge/

tangent-cone, independent-panel methods. For the

higher hypersonic Mach number regime, Newtonian

methods are used to compute surface-pressure

coefficients. Friction drag is computed using various

reference enthalpy methods (e.g. Eckert or van Driest).

The engine-off base drag is computed using empirical

relations for the high-subsonic/transonic/lowsupersonic speed range, and estimated at higher Mach

numbers assuming a 70-percent vacuum in the base

region (i.e., Cp_ = -1/MS).

Calculation of the aerothermal environment begins with

the flight condition along the vehicle trajectory. With

the flight condition specified (i.e., given the free-stream

Mach number, the angle of attack, and the free-stream
dynamic pressure), the boundary-layer edge conditions

are computed for each panel. For acreage-heating

calculations, the skin-friction coefficient is computed

using reference enthalpy methods noted above.

Calculations of local skin friction and convective heat

transfer film coefficient involve an iteration to balance

the convective heat flux, the radiative heat flux, and the

conductive heat flux into the vehicle. For vehicle

designs with exterior insulation, the thermal energy

conducted into the TPS is generally on the order of

5 percent of the total convective heating. Neglecting

this relatively small conduction term in the surface

energy balance results in a somewhat higher radiation

equilibrium surface temperature, which is used as a

design margin for the TPS. The reference Reynolds

number is then computed using the boundary-layer

edge velocity determined above and the running length

for that panel location. The local skin-friction

coefficients and local Stanton numbers are then

calculated using the Reynolds analogy and the

reference Prandtl number with the appropriate

formulation, depending on laminar or turbulent flow

conditions. Finally, the convective heating film

coefficient is computed from the Stanton number,

reference density, and edge velocity.

The local recovery enthalpy is determined using the

edge static enthalpy, the edge velocity, and the edge

Prandtl number, with the Prandti number correction

factor for laminar or fully turbulent flow, as appropilate. The convective heat-flux rate is then computed.

The radiative heat flux is computed by determining the

emissivity as a function of TPS material distribution

and the wall temperature. The iteration loop is repeated

until surface energy balance is achieved.

For blunt, leading-edge heating calculations, a modified

stagnation heating Faye-Ridell method _ is used. The

hot-wall heating rate is computed for off-stagnation

conditions using a modified Lee's method, t4 The

stagnation point is determined using a Newtonian

model with the unit surface normal co-linear with the

free-stream velocity vector. The actual wall temperature

computed for stagnation regions, or acreage areas, will

generally differ from the radiation equilibrium wall

temperature computed above. This conducting wall

temperature depends on the heating history along the

trajectory. Up to the peak-heating point, heat is

conducted into the TPS, and the resulting wall

temperature will be lower than the nonconducting wall

temperature. After peak heating, the TPS begins to

reject heat, with the resulting conductive heat flux back

to the surface resulting in a higher wall temperature

than the nonconducting value.

Propulsion

For the rocket engine, a simplified performance model

is used, with vacuum and sea-level-specific impulse as

inputs, along with vacuum thrust. Altitude performance

of the rocket is computed using ambient pressure

corrections to Isp.

LIFT

- BODY

- WING

DRAG

- INDUCED

- SKIN FRICTION

- BODY WAVE

- WING WAVE

- BLUNTNESS

-BASE

SUBSONIC TRANSONIC

Empirical f (AR, Match)

Low AR Theory

Empirical f (Mach)

Schlichting

Empirical

0 f (mean sweep, FR, Mach)

Empirical

0 f (sweep, T/C)

0 0

Empirical f (Mach)

SUPERSONIC HYPERSONIC

Tangent Wedge/Tangent Cone

Newtonian Mach >18

Linearized Supersonic

Tangent Wedge/Tangent Cone

Newtonian Mach >18

Reference Enthalpy

Eckert, Van Driest, Schlichting

ReThetaJM (Transition Criteria)

Tangent Wedge/Tangent Cone

Newtonian Mach >18

Prandtl Meyer Shock Expansion

LE Radius/Sweep & Newtonian

70% Vacuum

Figure 5. HAVOC aerodynamics.
For the ramjet, dual-mode and scramjet operation, a

nose-to-tail propulsion model uses the keel-line

geometry and quasi-one dimensional area distribution

of the combustor to compute the propulsion flow field.

This nose-to-tail propulsion model consists of a planar

inviscid 2-D real-gas, weak-wave flow code, coupled to

a 1-D subsonic/supersonic combustor analysis code, to

predict nose-to-tail flow-path characteristics. The

forebody boundary layer is computed using the inviscid

solution to provide edge conditions. Shear forces and

heat transfer are computed for the forebody/inlet using

reference enthalpy methods. Inlet unstart and self-start

boundaries are computed using computed flow-field

properties at the cowl-lip plane, and one-dimensional

isentropic flow between the cowl-lip and inlet throat.

Beginning at the inlet throat, the I-D combustor mass,

momentum, and energy equations with wall-skin

friction and heat transfer are solved (stepwise) through

the cumbustor. Multiple fuel-injector stations and preburning are provided for. Combustor efficiency (i.e.,

heat-release schedule) is computed as a function of

injector Mach number and injection angle, local

equivalence ratio, and axial distance from injector

station. Overall engine heat balance is computed using

an input combustor, skin-friction coefficient as a

function of free-stream Mach number.

In dual-mode operation, normal shock and thermal

choke plane locations are also computed. The nozzle

flow field is then computed from the combustor exit

solution using the real-gas, weak-wave 2-D code, which

includes nozzle and cowl flap geometry. Equilibrium or

frozen flow is computed in the nozzle region using the

mole fraction array at the combustor exit plane. Firstorder estimates of axial and normal forces and of

pitching moments are computed as a function of vehicle

geometry and flight conditions. Overall propulsion

system heat loads are then used to determine fuel inlet

total temperature or to compute required engine-cooling

equivalence ratio.

Trajectory

A 3-degree-of-freedom model is used to compute flight

trajectory, using the equations of motion of a pointmass aircraft moving relative to a rotating, spherical

earth. The trajectory flight path is computed using one

of two methods. _5First, an energy-state approximation

is used, coupled to an input schedule of Mach number

versus altitude. The second method is to specify a set of

control variables (consisting of angle-of-attack, bank

angle, throttle setting, and gimbal angle), and then to

solve the equations of motion, using a first-order Euler

6

integration (subject to a set of flight-path constraints,

including maximum load factor and/or surfacetemperature limits at a specified body location). For

either modeling methods, the trajectory can be

computed untrimmed or trimmed in pitch.

Structural and TPS Weights

As part of the aircraft structural weight, the items

estimated are: shell, walls, frames, tension ties,

spanwise beam, nonoptimum, and tank. The

nonoptimum weight accounts for non-calculable weight

items, such as fasteners, welds, cutout reinforcement,

surface attachments, nonuniform gage requirements,

and manufacturing constraints, and is a percentage of

the previous five weight items.

The tank item is an estimate of the weight of a discrete

tank or of bulkheads and other items necessary to

convert the body structure into an integral fuel tank. An

option for noncircular vehicles is an integral, pillowtank arrangement in which intersecting circular shells

are fitted within the vehicle outer mold line.

The body-structural-weight-estimation method in the

HAVOC is based on one-dimensional-beam theory

structural analysis, resulting in a weight estimate that is

directly driven by material properties, load conditions,

and vehicle size and shape, and is not confined to an

existing database. Since the analysis is done station-bystation, along the vehicle longitudinal axis, the

distribution of loads and vehicle geometry is accounted

for, providing an integrated weight that accounts for

local conditions.

The analysis begins with calculation of vehicle loads.

Three types of loads are considered--longitudinal

acceleration, tank pressure, and bending moment. Four

loading cases are computed--power-on/tanks full,

power-off/tanks full (abort condition), power-on/tanks

empty (MECO condition), and power-off/tanks empty

(entry condition).

It is assumed that structural materials exhibit

elastoplastic behavior. The values of properties used are

a reduced percentage of published, minimum values to

account for such effects as fatigue, stress corrosion,

creep, and thermal cycling; and thermal stresses that are

not modeled as cold structures are considered.

The body weight analysis described above has been

extensively correlated. It has been applied to existing

aircraft for the purpose of determining the nonoptimum

portion of structural weight. It has also been compared
withmanyotheranalysesofhypersonicaircraft.For

example, it wasappliedtothedesignsdevelopedbythe

fiveprimecontractorsforthefirstaircraftconcept

downselect in the NASP program, and agreement was

found to be generally excellent.

Estimation of TPS weight is based on transient, onedimensional heat-conduction analysis. The aerothermal

environments are computed as a function of time along

the flight trajectory, including the recovery enthalpy,

the enthalpy-based convective film heat transfer

coefficient, and the local surface pressure for each

surface panel. The one-dimensional materials stackup is

specified for each body point, with associated internal

boundary conditions (e.g., internal gas bulk temperature

and film coefficient), and internal radiation gap

conditions input as function of time. Soak-out boundary

conditions are also specified. Material thermoproperties are specified as a function of temperature and

pressure (if applicable). The energy balance at the

exterior surface is computed using a temperaturedependent emissivity for the corresponding TPS

material. The TPS sizing process then consists of

iterating on the required insulation thickness until all

interior bond-line temperature constraints are satisfied.

Minimum-gauge TPS thickness constraints are also

imposed.

The weight and volume of the vehicle subsystems are

computed using correlation equations developed ['or

space-launch and hyper-velocity vehicles. Correlation

parameters typically used related to overall vehicle

gross or empty weight, vehicle size, and mission-related

requirements. Correction factors for each subsystem

element are made available to match a specifically

known subsystem component weight. The accounted

subsystem elements are: flight controls and actuation

(electrical, hydraulic and pneumatic), power generation

and distribution, environment and life-support, thermal

control, and propulsion-feed systems.

Life-Cycle Costs

The life-cycle cost (LCC) is based on airframe and

engine development, vehicle acquisition, and

operational costs. Cost-estimating relations are based

on previous airplane programs, and these relations are

supplemented, when necessary, to account for the fact

that spaceplanes approach, rather than actually have,

aircraft-like operation. The following databases were

used: B-727, B-737, B-747, and B-757 (modified with

those for X- 15 and XB-70), engine (HyFAC Report_6),

launch vehicles (other than the Space Shuttle), and

operations (airline-like). Dry weights and vehicle speed

are largely used for estimating development and

acquisition costs. In the final analysis, the total cost of

space operations (including procuring and launching a

spacecraft) must be reduced.

Note that the cost-estimating relations for new vehicles,

based on new technologies and new operating

procedures, are likely to produce large uncertainties in

the estimated costs. However, the same cost-estimating

relations are used for all vehicles. The relative

differences are, therefore, much less uncertain than the

absolute costs.

Hot Versus Cold Structures

The launcher is modeled using 20 structural sections,

with skins, stringers, and frames sized for bending,

axial, and pressure loads. Insulation thickness is sized

for 16 zones, consisting of 20 perimeter points at 20

cross-sectional stations. Lobed tanks (fig. 6) are also

designed for bending, axial, and pressure loads.

Liquid hydrogen tanks are integral tanks having

intersecting lobes. These tanks are made of

graphite/epoxy structure with Z-stiffeners, and with

external Rohacell foam structure attached using

polyurethane glue (fig. 7). Toughened, unipiece fibrous

insulation (TUFI) ceramic tiles are attached to the

Rohacell foam tiles using RTV glue and a Kapton

vapor barrier. Liquid oxygen tanks on the orbiter are

Z-stiffened, aluminum-lithium tanks, with a similar

TPS. The leading edges of the spaceplane have the

same TPS as those assumed for the X-30 spaceplane.

Engine surface unit weights and seals, and systems

weights are also derived from those for X-30.

Figure 6. Integral tanks.
Tank

Tank Insulation

Structure ---__..

(If present)

TUF I Tile _.._

(Toughened

Unipiece

Fibrous Insulation)

Lower Vehicle

Surface 60°F

I

min profile

30 min profile

0 min profile

1500 °F

I

RTV glue

Gap or direct connect

Bond with RTV (Room

Temperature Valcunizing )

Tile thickness limits

structural temperature

to 400°F and minimizes

insulation weight and

boiloff in 16 zones

Temperature at time into mission

Figure 7. Thermal protection system (TPS).

Extensive development efforts during the 1960s

through the 1980s on integrated thermal structures (e.g.,

with TMC aeroshell), resulting in hot structures, have

not demonstrated an advantage over cool structures,

with external insulation, in terms of safety, robustness,

risk reduction, and simplicity. Hot structures are heavy,

complex, and expensive, particularly when the)' are

actively cooled.

Systems analyses for the launcher with Mach 5 staging

and 25,000-1b payload indicate that the best structural

approach for lightweight spaceplanes is using external

TPS on a cool structure (figs. 8 and 9). For these

analyses, structures are constructed using graphite/

epoxy (G/E) or aluminum alloys, and the external

insulation of these structures is tailorable advanced

blanket insulation (TABI) or TUFI. Cold, integral-tank

structures are 23 percent lighter than hot-shell

structures for hypersonic cruisers. _7The research,

technology, development, and evaluation (RTD&E)

cost estimates for G/E structures with external

insulation are significantly lower than the cost estimates

for hot structures (fig. 10).

The external insulation design for spaceplane is lighter

than those based on hot structures because of: ( I )

structures with low peak temperatures utilize more

efficient and less expensive materials, (2) when the

insulation is placed outside a structure, the usable

volume of the structure increases, (3) the TPS is less

complicated, and (4) insulation of cryogenic tanks

leads to a fewer requirements on TPS.

t:

Launcher Characteristics

[] .... I

[] .... I

Hol 81fuel G/1E÷TABI+Cp _ ÷ TUFI-12 G/E + TUFI-8 AL ÷ TUFNI

TPS Concept

Figure 8. A comparison of system weights for different

TPS concepts.

Installation of ceramic TPS on a cool structure, in turn,

does impose design and integration issues on the

airframe, including allowable strain constraints on the

substructure for direct bond, requirement for launcher

panels in nonconformal tank regions, and surface shear

and vibro-acoustic limitations associated with blankets.

Standoff metallic TPS present integration issues, in

terms of support structure, lower temperaturc limits,

8
face-sheet tmperaturegradient andthermalstress, and

creeplimitations. Ceramicsystemsaregenerallyighter

thanmetallicTPSasaresultofrelativelyhigh

attachment unitweightformetalliconcepts, thelatter

becominglighteronlyforhigh-integrated heatloads

whenexpensive, lightweight multilayerinsulation

isused.MetallicTPsystemsrequirecoatings, whichif

damagedleadtochemical reactionsatthemetallic

surfaceandtoacatastrophic failure.

Launcher Characteristics

=E

1.0 _

3:

o.s

0.o

HOt S_ct G/E÷TABI+CP G/E + TUFI-12 _ ÷ TUFI-8

TPS Concept

1.s t ...............................................................................................

AL + TUFI-8

Figure 9. A comparison of take-off gross weights with

different TPS concepts.

The X-33 was designed with cold structures. It has

load-carrying tanks and an external, leeward aeroshell

and windward frame structure that supports the

metallic, standoff TPS, and transmits aerodynamic

loads into the primary structure. The proposed TSTO

spaceplane has directly bonded TPS.

Optimum Staging Mach Number

A study is being conducted to assess the impact of the

staging Mach number on the size, weight, performance,

and life-cycle cost of the proposed TSTO spaceplane.

The LCC is the least, if the staging Mach number is

approximately Mach 10, as discussed below.

As the staging Mach number is increased, the weight of

the TPS increases on the launcher, increasing

significantly above Mach 10 (fig. 11), due to high,

convective heat loads associated with high-Mach and

high-dynamic pressure-flight conditions. Likewise, the

TPS fractions for the orbiter increase (but only

slightly), as it is carded on the launcher and entry heat

loads primarily determine TPS weights. Apart from the

harsher thermal environment at higher flight speeds, the

size of the launcher slightly increases, with increasing

fuel fractions, as the staging Mach number is increased.

Empty weights of both stages decrease as the staging

Mach number is increased, until Mach 12 (fig. 12).

Beyond Mach 12, the empty weight of the launcher

increases, driven by the increased TPS weight and the

increased structural weight as a result of higher,

required mission-fuel fractions.

LauncherCharacteristics

1.2

1.0

d0.8

LU

e_

k... 0.6

ee

0.4

0.2

0.0

Figure 10.

!il__?_ili_i_i i

HotSIbruct G_+TABPeCP G/E+TUFI-12 G/E+TUFb8 AL÷TUFI-4

TPS Concept

A comparison of RTD&E costs for different

TPS concepts.

As the staging Mach number is increased, the TOGW,

which is the takeoff gross weight of the launcher,

decreases until Mach 12 (fig. 13). The orbiter weight

decreases as less fuel and oxygen are required for

achieving reduced velocity increment to reach orbit

with the higher staging Mach number. The launcher has

to carry a lighter obiter, resulting in lower structural

weights for the launcher as the fuselage bending

moments are reduced. Thus, the total system weight is

reduced for the same payload or the payload weight is

increased for the same system weight, while the basic

physical characteristics of the spaceplane remains the

same as the staging Mach number is increased from

5 to 12.

When the Mach number is greater than 12, the required

engine-equivalence ratio increases above the

stoichiometric value to meet engine-cooling

requirements, with an associated reduction in the

engine-cycle, specific impulse. This leads to higher

mission-fuel fractions and, hence, higher closure gross

weights. Finally, staging at a higher Mach number

increases the downrange staging point, resulting in a
longercruise return to launch-site range requirement.

This further increases the required mission fuel fraction,

and pushes the vehicle closure point to higher gross and

empty weight values.

Mission: 25K Payload to 220nm Orbit @ 51.6oo

14.0

12.0 • -- -- Orbiter

"•10.0 ' ?

,_ 8.0 /

_ 6,0

4.0

launcher

J I

6 8 10

Staging Mach Numbe_

12

Figure 11. Sensitivity of TPS weights to staging

Mach number.

14

400

,,D

 ,o0

z_

;=

E

MJ

o 1oo

J¢

0

>

0

4

_"_" _ ,..,......__

Launcher

To_I //_

8 lo

Staging I_ch Number

I

12 14

Figure 12. Sensitivity of vehicle empty weight to staging

Mach number.

The RTD&E costs are driven primarily by subsystem

weight or, in the case of the airframe RTD&E costs, by

a dry weight-speed product. The acquisition cost is

primarily determined by dry weights, and operational

costs are largely determined by maintenance and

propellant costs. Maintenance costs are primarily a

function of subsystem dry weight and vehicle surface

area.

1,000

es 8OO

o

,pca

2

o

Q

.¢

0°

4

launcher

3

J

J

6 8 10 12 14

Staging Mach Number

Figure 13. Sensitivity of TOGW to staging

Mach number.

Figure 14 shows that the LCC decreases as the staging

Mach number is increased, from Mach 5 to 10. From

Mach 10 to 12, RTD&E costs of the launcher offset the

reduced empty weight trend, resulting in increasing

LCC. The cost of RTD&E and acquisition of scramjet

engines, operating above Mach 10, increases. Existing

ground-test facilities, and test techniques and subscale

flight tests, such as those planned for X-43, are

inadequate for developing scramjet engines above

Mach 8. New test facilities and techniques, and flight

tests with scramjet engines approaching those on

spaceplanes, are required. 10

Figure 15 shows the general configuration layout and

dimensions of the spaceplane for Mach 10 staging.

Table 1 presents some details of mass properties for the

launcher and the orbiter for Mach 5 and 10 staging

conditions. For the launcher, the weights for the

horizontal and vertical tails and for their TPS are bookkept under control surfaces. The payload for the

launcher consists of the full-up orbiter launch weight,

plus integration/separation-system increment. The

takeoff thrust-to-weight of the Mach 5 launcher is

0.516.

10
1.2

W

'_ 1.0

0

o

E

I_ 0.8

2

a.

(U

0,6

I--

0 0.4'

n-

-----( )_....

Launcher

F

Total

-- r

_ O_iter

,

8 10

Staging Mach Number

I

0,2'

4 6 12

Figure 14. Sensitivity of LCC to staging Mach number.

The trends presented in figures 11 through 13 would be

the same, even if some system weights were computed

differently. For example, the landing gear weight is

assumed to be approximately 1.5 percent of the TOGW,

If this weight were to be 2.0 percent of the TOGW, the

trends observed in these figures would not change.

A Prototype/Experimental Spaceplane with

Built-in Growth

A TSTO spaceplane with Mach 5 staging offers lower

performance at lower risk (because it avoids a

scramjet), while that with Mach 10 staging offers

higher performance at moderate risk (with a scramjet).

Essentially, the same airframes can be used for staging

from Mach 5 through 10, with modest TPS weight

difference.

A basic design of a Mach 5 staging launcher is slightly

modified to stage the orbiter at higher Mach numbers,

or a latmcher designed to stage at Mach 10 is used for

launching the orbiter at lower Mach numbers. Likewise,

an orbiter designed to stage at Mach 5 is slightly

modified to stage at higher Mach numbers, or it is

designed to stage at Mach 10, and used for staging at

lower Mach numbers. Different combinations of the

launcher and orbiter will lead to different payload

launch capabilities. The combination giving the lowest

LCC and highest growth potential requiring the least

modifications is the one developed.

>'1 I<_ Gross Wt. = 619151 Ibs

52'-4" Empty WIL= 251638 Ibs

_.t

30'-8" WG = 203792 Ibs

WE = 33235 Ibs

Figure 15. TSTO spaceplane, Ms,= 10: launcher and orbiter.

11
Table 1. Mass properties for Mach 5 and Mach 10 spaceplanes.

Item

Airframe

Wing and TPS

Fuselage

Fuselage TPS

Control Surfaces

Landing Gear

Propulsion

Ramjets

Rockets

Turbojets

Inlets

Fuel Systems

Fixed Equipment

Empty Weight

Consumables

Propellant

Payload

Gross Weight

Mach 5 Staging

Launcher, Ib

113,695

0

68,264

15,271

15,842

11,943

133,914

41,563

0

69,181

14,748

8,233

17,269

264,879

1,176

102,331

476,490

848,846

Orbiter, Ib

30,203

1,652

12,545

8,371

4,810

3,322

17,011

0

12,088

0

0

4,924

9,689

57,403

1,920

389,114

25,000

474,168

Mach 10 Staging

Launcher, lb

110,984

0

65,769

18,103

17,059

10,052

125,858

39,179

0

65,175

13,828

7,675

14,405

251,638

1,759

157,840

204,813

619,151

Orbiter, lb

18,832

621

7,588

5,621

3,758

1,816

9,296

0

6,605

0

0

2,691

4,535

33,235

1,954

143,603

25,000

203,792

As discussed in reference 10, the development of

scramjet propulsion requires tests of engines that are to

be used on spaceplanes, in the absence of at least two

(preferably three) appreciably different size subscale

engine tests.

The principal scramjet development challenge is in the

Mach 10 to 23+ range. Development of a prototype

spaceplane leading to a fleet of operational vehicles

requires: (1) a demonstration of net scramjet thrust

across the complete air-breathing hypersonic Mach

number range of interest, (2) validation of simulation

models and verification of simulation-design tools, and

(3) verification in an actual vehicle of the technologies

and systems required for such vehicles.

These observations suggest the following programmatic

philosophy:l° First, a TSTO spaceplane is designed for

Mach 10 staging to orbit with at least a 25,000-1b

payload to the design orbit (51.6 ° and 220 n. mi.).

Second, the spaceplane is used for space access by

staging at Mach 6. (The reason for choosing Mach 6

rather than Mach 5 for staging is explained later in this

section.) Third, a fully reusable rocket propulsion

system is developed. Fourth, the launcher is used for

demonstrating scramjet operations from Mach 6+ to 10.

Fifth, once the scramjet is perfected to the level

necessary and made operational, the spaceplane is used

for space access by staging at Mach 10. Sixth, further

development of scramjet technology at higher Mach

numbers is conducted on a new, developmental

second stage.

Following this philosophy, a P/X TSTO spaceplane is

recommended as a means of providing access to space,

for developing the operational hypersonic air-breathing

propulsion system, and for developing fully reusable

rocket-propulsion system, with significantly improved

mass fraction and margin. 1°Note that an aircraft may be

a fighter aircraft and at the same time be an attack

aircraft, e.g., F/A- 18. In principle, the same plane may

be a prototype plane and an experimental plane, and

could be called a growth plane.

Three orbiters are developed (table 2). The launcher

and the orbiters are reusable, piloted, and takeoff and

land horizontally. The airframe and the turbo/ramjet on

the launcher are fully reusable; the orbiter has fully

reusable airframe and has low-maintenance rocket

engines.

The launcher is initially designed to achieve Mach 10.

For example, the design of the launcher would include

the TPS thickness distribution required to accommodate

heat loads associated with Mach 10 operation and fuel

tank capacity for the Mach 10 mission. The launcher

airframe is a prototype. The launcher has a prototype

propulsion system to achieve Mach 6 and has a

12
demonstrator/experimental, dual-moderandscmmjet

enginefromMach6+to10.The flow path of the

experimental dual-mode ram/scramjet is the same as

that for the prototype ramjet. The upper Mach 10 limit

for the launcher is selected considering the staging

Mach nmnber trade study.

Table 2. Characteristics of TSTO P/X spaceplane.

Ve_c_

Launcher

Orbiter-R

Orbiter-E

Orbiter-A

Mach

Range

0 to 6

6+to 10

6 to orbit

9 to 24

9 to orbit

Propulsion

Turbo/ram (P)

Ram/scramjet (X)

Rocket (P)

Rocket(O)/Scramjet (X)

Scramjet(P)/Rocket (O)

Airframe

P

P

P

P

P: PROTOTYPE, X: EXPERIMENTAL, AND

O: OPERATIONAL

After an operational scramjet that can perform up to

Mach 10 is developed, a block change of the prototype

Mach 6 propulsion system and experimental Mach 10

engine to a prototype Mach 10, air-breathing propulsion

system leads to a prototype launcher for Mach I0

staging.

The three orbiters are: (1) orbiter-R with an all-rocket

propulsion cycle, (2) orbiter-E with a rocket/airbreathing propulsion system, and (3) orbiter-A with an

air-breathing/rocket system. Orbiter-R is a prototype

vehicle, designed to go to orbit. Orbiter-E and orbiter-A

are discussed later on.

Initially, orbiter-R is staged at Mach 6, because staging

at Mach 6 reduces LCC from that for Mach 5 staging,

while requiting little increase in the technology level

for the launcher. The value of initial staging Mach

number is chosen low to build-in payload growth to

Mach 10, with the same launcher and the same

orbiter-R.

The tanks on orbiter-R are sized for a Mach 6 launch.

Consequently, orbiter-R would have excess propellant

volume when launched at Mach 10, and may achieve

higher altitude orbits than the design orbit with lighter

payloads. At higher staging speeds, the tanks are

partially filled for achieving design orbit, so that the

eliminated propellant load is replaced by additional

payload weight within the existing payload bay.

Additionally, payloads with larger size could be

accommodated by rearrangement of the internal

propellant tank configuration.

Table 3 presents mass properties for the launcher

designed to stage the orbiter at Mach 10, but initially

stages the orbiter at Mach 6, and for the orbiter. The

Mach 6 staged orbiter is capable of carrying a 8,000-1b

payload.

Strategically, the next-generation transportation system

should be put into service as soon as possible, even

though, initially, it would transport payloads much

lighter than 25,000 lb. If, during the initial operations of

this commonality spaceplane (with extra margins in the

design and launching lightweight payloads), RTD&E

are undertaken in flight, an early IOC can be achieved.

As discussed later, the time saved in achieving early

IOC saves billions of dollars in total launch costs.

The launcher and orbiter-R can be developed with a

high level of confidence, since vast amounts of data,

information, and knowledge are available from lowspeed systems, ramjet systems, the space shuttle orbiter,

the NASP program, and other atmospheric-entry

vehicles. Although the space shuttle main engine

(SSME) requires a major maintenance/overhaul after

approximately four missions, data, information, and

knowledge are available for building low-maintenance

rocket engines that can be reused for approximately 15

missions without a major maintenance. (A few Russian

rocket engines are claimed to achieve majormaintenance-free operations equivalent to

approximately 10 missions.) The development of rocket

engines that can be reused for approximately 50

missions, without major maintenance, will require

technology efforts.

As discussed in reference 9, relevant technologies for

all subsystems are low-risk technologies. For example,

the ramjet was successfully ground tested to simulated

Mach 8 conditions in the 1960s.

Because a significant portion of the evidence for

establishing the credibility of the design would be direct

evidence, the level of confidence in the design of the

launcher would be quite high, to speeds of

approximately Mach 6. The quantity of this type of

evidence would decrease and the level of indirect

13
evidencewouldincrease, as Mach 10 is approached.

Also, the level of confidence in the design of orbiter-R,

with major-maintenance-free rocket engines for

approximately 15 missions, would be high.

Table 3. Mass properties for Mach 6 launch.

Item

Airframe

Wing and Tail

Fuselage

Fuselage TPS

Control Surfaces

Landing Gear

Propulsion

Scram jets

Rockets

Turbojets

Inlets

Fuel Systems

Fixed Equipment

Empty Weight

Useful/Consm.

Propellant

Payload

Gross Weight

Launcher Orbiter

M,, -- 10, lb M a = 6, lb

95,455

13,725

55,124

15,700

2,448

8,458

126,477

54,067

0

49,609

13,264

9,536

14,007

235,939

3,178

161,155

200,700

602,388

17,271

2,732

7,263

5,046

819

1,401

7,174

0

5,097

0

0

2,077

6,329

30,773

2,423

158,095

8,000

199,301

The experimental scramjet engine on the P/X

spaceplane would provide direct evidence in the Mach

6 to l0 range for developing prototype ram/scramjet

14

propulsion system. Initial flight experiences with

orbiter-R would lead to the development of lowmaintenance, 50-mission rocket engines.

Primarily, the integration of turbojet and ramjet

engines, ramjet-scramjet transition, vehicle

performance at low-hypersonic Mach numbers, stage

separation, and the rocket propulsion system (with a 50-

mission operability at low maintenance) are the risk

items. These are relatively minor risks compared to the

technical risk of developing a commercially viable,

SSTO-R vehicle for IOC in 2012.

Since the access to space study, s hypersonic

technologies have advanced and new hypersonic

technologies have been developed. Today, all required

technologies for developing the proposed TSTO

spaceplane, that would stage the orbiter at Mach 6 with

air-breathing propulsion on the launcher, are believed to

be at technology readiness level s (TRL) of at least 5. All

TRLs can be brought to level 6 within the next 3 years.

Performance at Mach 10 Staging

The flight trajectory of the spaceplane with a 25,000-1b

payload during ascent is a typical airplane/air-breathing

type trajectory with an initial, low-speed acceleration at

low flight-path angle (fig. 16). The transonic flight

regime is flown at an approximate dynamic pressure of

1,250 lb/ft 2 to reduce the induced aerodynamic drag at

the transonic minimum acceleration or pinch point. The

dynamic pressure is about 1,500 lb/ft 2 from Mach 2 to 4

for supersonic climb. Subsequently, the dynamic

pressure is gradually increased to approximately 2,100

lb/ft _during acceleration at hypersonic speeds.

A 2.0g pull-up maneuver is initiated at Mach 10 and at

an altitude of 95,000 fi, to lower dynamic pressures and

achieve the optimal launching flight path angle for the

orbiter vehicle (fig. 17). The MECO is at Mach 10.4

and at an altitude of 106,000 ft. The orbiter is launched

at an altitude of approximately 140,000 ft. and the

dynamic pressure of 300 lb/ft 2.After second-stage

separation, the launcher continues its unpowered ascent

and initiates a 66.4 ° bank to begin the return-to-launchsite maneuver. Maximum attained altitude is

approximately 200,000 ft. During entry, a turn to the

cruise-back heading angle is completed and the

launcher cruises at maximum specific-range Mach

number (2.0), using the low-speed turbojets to return to

its launch site. The orbiter rapidly climbs and

accelerates to approximately Mach 27.5, when it

reaches the MECO condition and then coasts to the

desired orbit. Figure 17 also shows the orbiter entry
flight.It useslow,initial-entryflight-pathangleto

ensurethatheatloadsarewithinthedesignmargin.

First Stage Vehicle

100,000 - 3,000

80,000 _

¢_. 60,000 ,J _x_ _ _

._- _ _'_

--" 40,000 L 0

-,,ooo

20,000 P'

o /<_ o

0 2 4 6 8 10

Mach Number

Figure 16. Trajectory and dynamic pressure

experienced by the launcher.

A possible abort mission for the spaceplane consists of

taking off, climbing, accelerating to the launch point,

performing the pull-up maneuver, and then aborting the

launch and returning to the takeoff point (fig. 18).

When a decision is made to abort after MECO, and just

before reaching the launch point, the spaceplane begins

a roll to 66.4, reaches maximum altitude of 206,000 ft

at Mach 8.76, and then begins to descend. During

atmospheric entry, the spaceplane experiences

maximum dynamic pressure of 730 lb/ft 2at Mach 7.5

and at an altitude of 106,000 ft. It reaches the maximum

down range at Mach 6.4 and an altitude of 111,000 ft.

After completing the U-turn, at normal loads less than

2.5g, it is travelling at Mach 2 and is at an altitude of

75,000 ft. At this point, it begins turbojet power-on

cruise back to the takeoff point at the best specific

range Mach number. During the abort phase, the

launcher continues to carry the orbiter at its full launch

weight.

If a launcher engine shuts down at any point during the

trajectory, the spaceplane can return safely to the

airport of origin the same as an airplane would. The

launch-abort mission impacts the sizing of the thermal

protection system and fuel load on the launcher•

The bending moment distributions on the launcher are

shown in figure 19 for power-on/tanks full, poweroff/tanks full (abort condition), power-on/tanks empty

(MECO condition), and power-off/tanks empty (entry

condition). For all loading conditions, the launcher is

carrying the orbiter at its launch weight. Bending

moments are distributed fairly symmetrically around

the center of mass for all conditions•

¢:

"O

-i

,i..,

<

Mission: 220nm orbit at 51.6

450,000

400,000

350,000 Orbiter Ascent J

300,000 f -'_

f J

2so,ooo / _

/ j ....-"

200,000 "LBuncher -- Ib_ f" _ _ vr_

Descent ! _ / _ Orbiter Entr ,-

IS0,000 ., _ I I

_ _ Launcher Pull-up

100,000 _" / [ I

S0,000 --_j '_ --" Launcher Ascent

0

0 5 10 15 :'0 2-_ 30

Mech Number

Figure 17. Launcher and orbiter trajectories.

Figures 20 and 21 show temperature time histories at

specific, in-depth locations on the launcher and orbiter•

Temperature histories at the surface, at the bond line

between the external insulation and the internal

cryogenic insulation, and at the interior surface of the

internal cryogenic structures are plotted. The peaks in

profiles for external surfaces are associated with peak

heating environments encountered by these structures.

The bond line temperature remains below 350 F, a

design constraint. The resulting internal temperatures

on the launcher's cryo-structures are well below

material allowable limits.

Internal Rate of Return

"Fable 4 compares a spaceplane designed to stage at

Mach 5 with a 25,000-1b payload on the orbiter, with a

spaceplane whose orbiter is launched either at Mach 5

or 10. The TOGW of the latter spaceplane (the

commonality spaceplane) and the weight of its orbiter

are the same, whether the payload weight is 8,000 Ib

launched at Mach 5 or 41,000 Ib launched at Mach 10.

The launch cost per pound of payload, for launching a

payload weighing 25,000 Ib, with the commonality

spaceplane is approximately 0.92 percent of that for the

baseline spaceplane (table 4).

15
E

f..

O

Z

v

O1

cn"

8OO

700

600

5O0

400

300

200

100

Mission: 220 nm at 51.6 °

ange

= 6.4/111K

Max Entry Dynamic Press:

MaclVAlt = 7,5/106K

'.MaxAltitude: MacWAIt = 8.76/206K

0 50 100 150 200 250 300 350 400 450 500

Range (East), nm

Figure 18. Spaceplane launch, abort, and recovery ground track.

First Stage Bending Moment vs. Body Station

j_ 40,000,000

"T

¢:

C

30,000,000

E

O

O1

.--- 2o,oo0,ooo,

"0

C

¢Q

01

lo,ooo,ooo,

,,-I

I,i.

0

0

_._ _....... _ . j _ Tanks Full/Power On

L • _. : Tanks Full/Power Off

i i _ i! _ TanksEmpty/powerOn

:: _ _:: Tanks Empty/Power Off

.......... i......... I......... i'";Ir_V ....... _.......... II.............. I.......... i........ i_

50 100 150 200 250 300

Body Station, ft.

Figure 19. Launcher bending moments.

]6
20MinHoldTime 20Min + CapturedFlight

1,500 2,500

,m

erM.

Q.

E

1,000

500

0

2,000

,m

• 1,500

= \

E

• External _ J

.¢:

u.m 1,000 -_,

_

,oo0 =--_

Internal Bond _ .._ _

Internal -500

-500 -1,000

1,000 2,000 3,000 0 1,250 2,500 3,750 5,000

Time, sec Time, sec

Figure 20. Launcher surface and bond-line temperature

histories at 25 ft from the nose on the bottom surface.

Figure 21. Orbiter surface and bond-line temperature

histories at 25 ft from the nose on the bottom surface.

The acquisition cost is for a fleet of five operational

launchers and seven operational orbiters plus two groundtest vehicles, one for the launcher and one for the orbiter.

Thirty-nine flights per year are operated for 22 years. Cost

estimates are in 1992 U.S. dollars.

The LCC for each of the three systems (all-rocket SSTO,

air-breathing/rocket SSTO, and air-breathing plus allrocket TSTO) considered under option 3 of the NASA

access-to-space study x are almost the same as those for

the other two systems (fig. 22).

The TSTO spaceplane is a low-risk concept, requiring

low-risk technology, compared to the SSTO rocket

vehicle or the SSTO air-breathing/rocket vehicle. The

sensitivity of the SSTO rocket vehicle to small changes in

dry weight and in specific impulse is significantly larger

than that of the TSTO spaceplane (fig. 23), because the

SSTO rocket has a very high fuel fraction requirement,

approximately 90 percent.

Table 4. Weight and LCC comparisons of

TSTO spaceplanes.

Item Baseline (B) Commonality

Payload, Ib 25K @ M_, = 5

Weight, Ib

Launcher Empty

Orbiter Empty

Total Empty Wt

Takeoff Gross Wt

Cost

(1992 dollars)

RTD&E

Acquisition (A)

Operations (O)

TOTAL (T)

Launch Cost/lb (L)

265K

57K

322K

849K

SBrI1)&E

$BA

$Bo

$Bi

25K lb @ $Lt_

8K @ M,,=5

41K @ M,,=I0

252K

33K

285K

619K

_B ria}&l.,

$(B a- 10")

$Bo

$(B,-10 ')

8K lb @ $(2.92xL,)

25K Ib @ $(0.92xL,)

41K lb @ $(0.58xL n)

17
1.2

1,0

0.8

0.6

0.4

0.2

0,0

RTO&E

_.quisition

OP_rations

Total Life-Cycle Costs

ssTo(m SSTO(,Ms.) TSTO(m) c(..mo_,_t_.

(M6-MIO)

Access-to-Space Concepts

Figure 22. LCC for the access-to-space study vehicles.

success; that is, it should be economically viable. The

RTD&E and acquisition costs of the new system, added to

the overall costs of continued operation of the space

shuttle and the existing fleet of expendable launch

vehicles, leads to the increase in the launch vehicle total

cost (fig. 24). When a new fleet of reusable vehicles

replaces a present fleet, the new fleet would greatly

reduce operating costs.

1.50-

O

O 1.25-

_ 1.00.

I£

-,!

c

c 0.75.

<

tu

0.50.

"_ 0.25 -

0,00

1995 2005 2035

I_ TSTOM=5,1RR=14% =17%

i i

2015 2025

Year

°_

C3

500,000

450,000

400,000

350,000

300,000

250,000

200000, -

150,000 -

100,000 -

50,000 -

O4

0.9

SSTO(R): Dry Wt

__ Dry Wt

\

SSTO(R): Isp

I I I

0,95 1.05 1.1

Dry Weight and Isp multiplier

I

1.15

Figure 23. Sensitivities of SSTO rocket-powered vehicle

and TSTO spaceplane. 9

A new launch vehicle meeting commercial requirements

and interests must create profit to be a commercial

Figure 24. Funding streams over time, with internal rate of

return (adapted from reference 9).

The Mach 5 to 10 commonality TSTO spaceplane

substantially reduces the need for experimental vehicles.

Additionally, the commonality spaceplane, with early

operations at Mst = 5, and with subsequent operations at

M_, = 10 l-year later, saves 10s of billions of dollars over

a 5-year period. 9The TSTO commonality spaceplane

achieves operational status much earlier than the SSTO

rocket vehicle and the TSTO Mach 5 staging spaceplane,

because the latter vehicles are required to have a 5-year

technology program to reduce risk with flight

demonstrations. In the case of the commonality

spaceplane, technology development and risk reduction

are carried out in flight, while light-weight payloads are

launched to reduce total launch costs.

The SSTO rocket system was estimated to provide

I I-percent internal rate of return (IRR). The commonality

spaceplane offers a 17-percent IRR. This rate is more than

that provided by the SSTO rocket and the TSTO vehicle

with Mach 5 staging. However, 17-percent IRR is not

high enough for an economically viable launch system.

18
Todramaticallyincreasethesafetyandreliability,

drasticallyowertheLCCofaccesstospace, andto

convergenoncommercial/commercial requirements and

interestsforspacelaunchmaybedifficult.Asobservedin

reference9,a22-percent returnoninvestment maybe

feasible, if theGovernment fundsthefirst2yearsofthe

development program.

Further Growth Potentials 1°

The reusable and operationally flexible launcher provides

vital access-to-space launch and hypersonic flight test

services capability. However, the orbiter-R provides

short-term economical benefits by achieving orbit for

space missions, while orbiter-E (table 2) is used for

further scramjet development. This orbiter serves as a

testbed for conducting experiments and developments at

high dynamic and heating loads, such as those related to

full-scale structural panels and components, including

scramjet engines. Orbiter-E is utilized as the X-7 and X15 planes were.

The Mach 5 to 24 range is divided into three ranges, low-

(Mach 5+ to 10), moderate- (Mach 10+ to 18), and highMach number (Mach 18+ to 24) ranges. This division

facilitates testing of the scramjet operation over the lowhypersonic Mach range, with the launcher and the

incremental development of air-breathing propulsion with

orbiter-E at moderate- and high-hypersonic Mach

numbers. This divide-and-conquer philosophy

significantly reduces development and flight-test risks.

Orbiter-E is designed to go from Mach 9+ to 24, and is

primarily a rocket-powered vehicle with only one

replaceable air-breathing engine. The orbiter is designed

to fly, when required, selected parts of orbiter-A's airbreathing trajectory. The configurations of orbiter-E and

orbiter-A may be different from that for orbiter-R.

The development of the propulsion system with orbiter-E

is accomplished in three steps of increasing technological

challenges: from Mach 9 to 12, Mach 12+ to 18, and

Mach 18+ to 24. Once this system is developed in the

Mach 9 to 12 range, the prototype orbiter-A is built.

When the scramjet is developed in the Mach 12+ to 18

range, the air-breathing range of orbiter-A can be

extended, or the development of a prototype SSTO

spaceplane with air-breathing/rocket propulsion can be

undertaken.

Orbiter-E is not built until the hypersonic propulsion

system performs satisfactorily in the launcher, up to Mach

10, and is well understood. While the launcher and

orbiter-R are designed, built, tested, evaluated, and made

operational, a program is carried out for advancing the

hypersonic facility capability and for improving

appropriate, nonintrusive, flow-diagnostic technology

applicable to the hypersonic environment. Flight-test data

from flights of the launcher in the low-hypersonic Mach

number range would improve computational-design

technology and calibrate ground-test data. These advances

and enhancements would help in the design of an

experimental air-breathing propulsion flow path for

orbiter-E, with a high level of confidence in its design.

Flight tests of this truly experimental vehicle in the

moderate- and high-Mach-number ranges and the

advances in hypersonic facilities would result in a high

level of confidence in the design of orbiter-A.

Rocket and air-breathing propulsion options are pursued.

These propulsion systems and the proposed vehicles open

up, for example, the following further growth potentials

and multiple avenues, any one of which may be pursued

with a high level of confidence: (1) replacement of

orbiter-R with orbiter-A, (2) development of an airbreathing/rocket SSTO vehicle, (3) development of an allrocket SSTO vehicle, (4) development of a hypersonic

cruiser, (5) development of an unpiloted orbiter, (6)

development of an expendable, unpiloted orbiter for highenergy orbits, and (7) development of reusable orbiters

for specific payloads.

Concluding Remarks

The foremost objective is to greatly reduce cost and

substantially improve safety and reliability for access to

space. In the near term, this objective can be best met

using a TSTO concept with a hypersonic air-breathing

propulsion system on the first stage.

Spaceplanes with low-risk technologies and built-in

growth potential and aircraft-like operations provide the

most return on investment. The conceptual TSTO P/Xspaceplane is closest to meeting economic

requirements--a 20-percent rate of return on investment.

This spaceplane significantly reduces risk, increases

margin, and maintains the SSTO option.

The P/X-spaceplane launcher assists the development of

hypersonic air-breathing propulsion and meets the desired

access-to-space requirements near-term. The orbiter

assists in the development of long-life, low-maintenance

rocket engines. This strategy offers a number of

advantages; is technology-driven, opens up multiple

future avenues, provides short-term benefits, has built-in

growth potential, and is achievable.

19
TheTSTOspaceplane, withair-breathingpropulsionon

thefirststage,appearstobethecorrectchoiceforthe

development ofhenext-generation, reusableaunch

vehicles. Thisconcept couldbeaseriouscandidatefor

IOCincirca2012,withMach6staging. TheTRLsare

believedtobeatorabovelevel5forthisspaceplane. The

prototypespaceplane withMach10stagingcouldbe

operational incirca2014.

References

1. Augenstein, B. W., Harris, E. D., with Aroesty, J., et

al., "The National Aerospace Plane (NASP):

Development Issues for the Follow-on Vehicle,"

Executive Summary, RAND, R3878/1-AF, 1993.

2. Hallion, R. P. and Young, J. O., "Space Shuttle:

Fulfillment of a Dream," The Hypersonic

Revolution, Vol. II, edited by R. P. Hallion,

Special Staff Office, Aeronautical Systems

Division, Wright-Patterson Air Force Base,

Ohio, 1987.

3. "National Aero-Space Plane, Restructuring Future

Research and Development Efforts," GAO,

NSIAD-93-71, U.S. Government Accounting

Office, 1993.

4. Rich, B. R. and Janos, L., Skunk Works: A Personal

Memoir of My Years at Lockheed, Little Brown

& Co., Boston, MA, 1994.

5. Domheim, M. A., "Engineers Anticipated X-33 Tank

Failure," Aviation Week and Space Technology,

15 November 1999.

6. "Final Report of the X-33 Liquid Hydrogen Tank

Test Investigation Team," NASA George C.

Marshall Space Flight Center, May 2000.

7. Bekey, I, "Forget About Shuttle Upgrades, Go With

RLVs," Aviation Week & Space Technology, 24

April 2000.

8. "Access to Space, Advanced Technology Team Final

Report," Vol. 1, Executive Summary, National

Aeronautics and Space Administration, July

1993.

9. Gregory, T., Bowles, J., and Ardema, M., "TwoStage-To-Orbit Air-breathing and Rocket System

for Low Risk, Affordable Access to Space," SAE

1994 Transactions, Journal of Aerospax'e,

Section 1, Vol. 103, 1995, pp. 189-201; also

SAE Paper 94-1168, April 1994.

10. Mehta, U. B., "Strategy for Developing AirBreathing Aerospace Plane," Journal of Aircraft,

Vol. 33, No. 2, March-April 1996. Errata in Vol.

33, No. 4, July-August 1996, p. 840. (This paper

is an updated and shortened version of "AirBreathing Aerospace Plane Development

Essential: Hypersonic Propulsion Flight Tests

(Invited)," Proceedings of the 2nd European

Symposium on Aerothermodynamics for Space

Vehicles, ESTEC, Noordwijk, The Netherlands,

November 21-23, 1994, ESA SP-367, February

19953

11. Morris, R. E. and William, N. B., "A Study of

Advanced Air-breathing Launch Vehicles With

Cruise Capability," Vols. I-VI, NASA CR

73194-73199, 1968.

12. Novichov, N., "At Hypersonic Speeds," FASTCID(RS)T-0972-92, Foreign Aerospace Science

and Technology Center, WPAFB, Ohio,

Translated from Ekho Planety, Vol. 42, No. 237,

17-23 October 1992.

13. Faye, J. A. and Riddell, F. R., "Theory of Stagnation

Point Heat Transfer in Dissociated Air," Journal

of Aerospace Sciences, Vol. 25, No. 121, 1958,

pp. 54-67.

14. Lee, L., "Laminar Heat Transfer Over Blunt Nosed

Bodies at Hypersonic Flight Speeds," Jet

Propulsion, Vol. 26, No. 4, April 1956.

15. Windhorst, R., Ardema, M., and Bowles, J.,

"Minimum Heating Entry Trajectories for

Reusable Launch Vehicles," Journal of

Spacecraft and Rockets, Vol. 35, No. 5,

September-October, 1998, pp. 672-682.

16. "Hypersonic Research Facilities (HyFAC) Study,"

Vol. V, Limited Rights Data, NASA CR 114330,

1970.

17. Private Communication, Integrated System Analysis

Team (ISAT), NASA Langley Research Center,

2000.

20

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February 2001 ] Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

A Two-Stage-to-Orbit Spaceplane Concept with Growth Potential

6. AUTHOR(S)

Unmeel B. Mehta and Jeffrey V. Bowles

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

Ames Research Center

Moffett Field, CA 94035- 1000

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

713-74

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AGENCY REPORT NUMBER

NASA/TM-2001-209620

1_, .SUPPLEMENTARY NOTES.; ..........

ius paper is presentect as AIAA Vaper 2OO1-179_ at the AIAA/NAS-NASDA-ISAS l 0th International Space Planes and

Hypersonic Systems and Technologies Conference, 24-27 April 2001, Kyoto, Japan. Point of Contact: Unmeel B. Mehta,

Ames Research Center, MS 229-3, Moffett Field, CA 94035-1000 (650) 604-6744

12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified -- Unlimited

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Availability: NASA CASI (301) 621-0390

13. ABSTRACT (Maximum 200 words)

A two-stage-to-orbit (TSTO) spaceplane concept developed in 1993 is revisited, and new information is

provided to assist in the development of the next-generation space transportation vehicles. The design

philosophy, TSTO spaceplane concept, and the design method are briefly described. A trade study between

cold and hot structures leads to the choice of cold structures with external thermal protection systems. The

optimal Mach number for staging the second stage of the TSTO spaceplane (with air-breathing propulsion on

the first stage) is 10, based on life-cycle cost analysis. The performance and specification of a prototype/

experimental (P/X) TSTO spaceplane with a turbo/ram/scramjet propulsion system and built-in growth

potential are presented and discussed. The internal rate of return on investment is the highest for the proposed TSTO spaceplane, vis-a-vis a single-stage-to-orbit (SSTO) rocket vehicle and a TSTO spaceplane

without built-in growth. Additional growth potentials for the proposed spaceplane are suggested. This

spaceplane can substantially decrease access-to-space cost and risk, and increase safety and reliability in the

near term. It can be a serious candidate for the next-generation space transportation system.

14. SUBJECT TERMS

Access to Space, Space Transportation, Spaceplane, Single-Stage-to-Orbit,

Two-Stage-to-Orbit, Air-breathing Propulsion

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