The performance goals of the NASP program require an aero-propulsion system with a high effective specific impulse. In order to achieve these goals, the high potential performance of air-breathing engines must be achieved over a very wide Mach number operating range. This, in turn, demands high component performance and involves many important technical issues which must be resolved. Scramjet Propulsion Technology is divided into five major areas: (1) inlets, (2) combustors, (3) nozzles, (4) component integration, and (5) test facilities. A status report covering the five areas is presented

Powell, William E.

English

NASA Technical Reports Server

N92-21534
NASP X-30 PROPULSION TECHNOLOGY STATUS
William E. Powell
Deputy. Director Systems Application
NASP JPO (NASA)
WPAFB Dayton, Ohio
Successful development of the NASP demands a propulsion system
which operates efficiently across the entire NASP operational flight envelope and at
speeds ranging from the takeoff to near-orbital velocity. To meet this challenge,
research is being conducted to .develop specific air-breattti_ng engine designs which
exhibit high effective specific impulse using combined subsonic-supersorfic-
combustion ramjet/scramjet propulsion concepts. Scramjet engine performance
critically depends upon effective, synergistic integration of new propulsion
technologies with the basic NASP airframe (see Figure 8-1).
New Matenaltar_ IStructures
I
Fiu_
/
l:igure 8-i. The Propulsion Challenge
The performance goals of the NASP program require an aero-propulsion
system with a high effective spedfic impulse. In order to achieve these goals, the
high potential performance of air-breathing engines must be achieved over a very
wide Mach number operating range. This, in turn, demands high component
performance and involves many important technical issues which must be
resolved.
22-I
https://ntrs.nasa.gov/search.jsp?R=19920012291 2020-03-17T13:09:57+00:00Z
Scramjet Propulsion Technology is divided into five major areas: (I)
inlets,(2)combustors, (3)nozzles, (4) component integration,and (5)testfacilities.
Criticalareas of focus for the component areas (inlets,combustors, and nozzles) are
the resolutionof key technicalissues,development of a high Mach number design
methodology, and establishment of a high Mach number performance data base that
will meet the challenging goals of the high performance and minimum weight
engine required forNASF. In component integration,integratedmodels of selected
component designs must be tested in order to resolve component integration
problems and to evaluate overallengine performance. Test facilitiesare required
(1)to provide Mach 5-8 testcapabilitiesof sufficientscale in order to conduct and
support the engine contractors'propulsion module testsand (2)to provide very
high Mach number simulations for smaller scalecomponent tests.
The scramjet inlet technology area addresses the key issues of inlet
"contraction ratio, inlet efficiency and air capture, boundary-layer effects and
simulation, shock boundary-layer interactions, and real-gas effects. The waves in
the internal portion of a hypersonic inlet tend to coalesce into a strong shock giving
r/se to a large adverse pressure gradient. Increasing the contraction ratio aggravates
the problem, thereby finally limiting the allowable compression ratio before
massive separation occurs. Relatively long forebodies are required to minimize
shock losses at high Mach numbers. Consequently, the boundary layer tends to
become relatively thick. The airframe shape and type of profile can have a
significant impact on inlet performance and its operating characteristics. Also, at
very high Mach numbers, the effect of 02 vibration can become important. Wave
structure of any given geometry is unique, and important inlet characteristics, such
as air capture, are difficult to match unless properly simulated. Combined analytical
and experimental efforts will provide answers to these issues, as well as develop the
methodology to design, test, analyze, and evaluate high performance hypersonic
inlets. Tests of small aerodynamic models will be conducted over a wide Mach
number range, including both wind tunnels and shock tunnels, and will be
complemented with applied computational fluid dynamics.
Hypersonic vehiclestend to utilizetheirlong forebodies as part of the inlet
compression process. This resultsin forebody boundary layersbeing ingested into
the propulsion system. In most cases,the complete forebody-inletsystem isdifficult
to model in a propulsion system test. Therefore, a technique to generate thick
boundary layersin supersonic flow must be developed with the proper momentum
defect distribution.
22-2
Studies in the scramjet mixing, area address the key issues of penetration,
wall and strut injection, supersonic shear layer mixing, and mixing augmentation
techniques. Experimental programs are underway to investigate shear layer mixing
and hypermixing concepts and to compare these results with CFD codes using
modified turbulence models. Several mixing augmentation techniques, including
lon_tudinal vorticity production and shock interactions, will be investigated
through university grants using the NASA Langley Mach 6 high Reynolds number
tunnel.
Shear flow developmentand mixing characteristics of noncircular nozzles
were investigated and compared to a circular jet over a range of Mach numbers at
the Naval Weapons Center (NWC), China Lake, California. Hot wire
measurements and sch/ieren photography were obtained. The superior mixing
characteristics of elliptic and rectangular jets relative to the circular jet, which were
known to exist for subsonic jets, were also found in the transonic jet and were
further augmented by the shock structures of the supersonic under-expanded jet.
Areas to be investigated in hypersonic mixing are effects of incoming
boundary-layer turbulence, longitudinal vorticity production, surface distortion, and
shock enhancement.
The scramjet combustor technology study area addresses the key issues of
film cooling/skin friction, ignition enhancement/flameholding, combustor
performance, diagnostics, and effects of initial conditions. At high flight Mach
numbers, protection of the combustor wall is of paramount importance due to the
extremely high enthaIpies of the incoming flow. Likewise, momentum of the fuel
is a major factor, and coax/al injection is requ/red for most fuel to maximize thrust.
Film cooling offers the possibility of simultaneously protecting the wall from
excessive heat flux and reducing wall shear. However, coaxial injection is not
conducive to rapid mixing. Measurements are not only more difficult to make, but
they must be more extensive than in a subsonic combustor since in supersonic
combustion there is no defined sonic point and exit property profiles are generally
nonuniform. Therefore, the entire combustor exit flow field must be measured to
accurately assess combustor performance and to provide initial conditions for
nozzle flow analysis. Combined analytical and experimental efforts, supplemented
by university grants, will clarify these key issues and provide sufficient
understanding to design a supersonic combustor capable of operating over a wide
Mach number range. New instrumentation techniques and laser diagnostics will
provide detailed flow-field measurements with which to calibrate computational
codes.
22-3
wThe scramjet kinetics study area addresses the issues of chemical kinetics,
reaction rate constants, and enhancement techniques for the three-body
recombination reaction. A chemical kinetic data base is being acquired for reliable
computer simulation of hydrogen/air supersonic combustion and for tests
performed in facilities using vitiated air. A shock tube and high temperature
kinetics cell, along with computational chemistry methods, are being u_lized to
obtain the critical rate constants at required accuracy over a wide range of
• temperatures. Identification of chemical additives that can speed up the exothermic
combining of radical species and experimental evaluation of their effectiveness will
be accomplished.
A sensitivity analysis of the hydrogen and air chemical reaction model
was performed by Los AJamos Natior_d Laboratory to identify which specific
reactions are the key rate-limiting steps in the heat release mechanism under
conditions relevant to scramjet propulsion.
The scramjet nozzle technology area addresses the key issues of
nonequilibrium thermochemical effects, fluid dynamic losses, thrust vector control,
and entrance profile effects. A major thrust loss mechanism in supersonic nozzles
at high Mach numbers is the thermochemical energy retained by dissociated species
when subjected to a rapid expansion process. Other mechanisms which lead to large
losses include wall skin friction and heat transfer, divergence, and internal
compression waves generated by nonuniform entrance conditions. Combined
analytical and experimental efforts will provide answers to these issues and
demonstrate internal nozzle performance, as well as develop a data base for flight
Mach numbers over a wide range of Mach numbers using both steady state and
pulse facilities.
The scramjet component integration technology area addresses the key
issues of combustor/inlet interaction, forebody effects on performance, and
combustor flow profile/nozzle performance. Flow profiles (including the nature of
the boundary layer) coming from one component will affect the performance of
subsequent components. For airframe-integrated scramjets, it is especially
important to investigate the effects of a simulated forebody flow on the performance
of the engine module. Combined analytical and experimental efforts will help
answer these issues, as well as develop a broad scramjet data base over a wide Mach
number range. Both vitiated and arc-heated freejet NASA Langley scramjet
facilities and the Calspan 96-inch shock tunnel will be utilized in establishing early
scramjet engine performance levels and resolve any key integration issues.
22-4
NASP IMPACT ON SSTO AEROSPIKE
ii
• NASP Risk Closure
ExWml Rocket
• NA8_ t.ln_r Rocket
..... ! |
• HASP Rlsk Closure
Exlemal Roc_t
• Fanned Rale_! Prog
• NASP CFD
• ALS
+. Lllerlgnll|on
• NASP Risk Closme
External Rocket
• ALS
• NASP Soa._et Fled ISystem
Qblectlve:
• Design, Build, Test X-30 Engine
Components to Demonstrate
Technology
- CFD Codes to Predict Inlet
Mass Capture, Combustion
Efficiency
• Revitalized National High-Speed
Propulsion Test Facilities
• Extensive Scremjet Data Base
• High Conductivity Materials for Heat
Exchangers
• Advanced 3D CFD Propulsion Codes
with Accurate Physical Modeltng for
Mixing, Combustion
Summary
Execution : NASP JPO, Contractor, GWPs
Funding : PE 63269F end NASA b
Fund+ng($M}" m m, .+ ,e_ Jog us m m m m mlln m m
T_-_"-_ Pflor 91 92 93i 94 95 96 97 96199 00 01 02 TOlal
X40 F._
EnOm Struct_w
MlWnll
APTU
FI_.t "real
val_lmon
(11 (2) I1(3) (4l
(5) (e)
(7)
,I
* Phase 3 Funding Estinmte Provided by Air Staff
"Actual Program Funding Requirement Due 2nd OTR FYO2
Milestones: +.
1. Concept Selection (4/91)
2. Size Freeze (2/92)
3. Technology Freeze Date (1/94)
4. Engine Delivery (4/97)
5. Material Flight Engine 11 Selecllon (1/93)
6. Structure Component Tests (3/93)
7. APTU Ram/Scramjet Flowpath Test Facility - FY96
22-5
• Provlds Propulsion Technology
and Development Test
- Develop Combustor Concepts
- Develop Integrated Engine
Configurations
• Enabling Technology for Wide
Range of Revolutionary Mission
Concepts
• Free World's Largest Hypersonic
Engine Test Capability
• Complete Testing Capability for
Airbreathing Engines up to Mech 8
• Full Range of Component Test Capability
Propulsion Test Facilities
Execution : NASP JPO, Contractor, GWPs
Funding : PE 83269F and NASA
91 92 93 94 95 96 97
(ETF)
sm_ smm
ASTF
Am'U
Oewl (1_
TeN
v_klmWn
• Phase 3 Funding EsUmate Provided by Air Stuff
• Actual Program Funding Requirement Due 2nd QTR FY92
1. Subecale High Match Combueter Development Facilities-
40 FY92
2. Static Test Stands - FY96
3. ASTF Syslem Test Facility. FY96
4. APTU Rem/ScramJet Flowpeth Tesl Facility - FY96
5. Component Test Fmcllltles. FY94
6. Full Scale Shock Tunnel for Combustor Development
7. LBRC 8' Hl"r Upgrades FY93
• Develop and Demonstrate
Hypersonic Airbraathtng
Propulsion Systems .
- Innovative Engine $iruclum
Concepts
- Large Scale ScramJet Data Blae
• High Speclflc Impluse Propulalon
Systems
. High Temp. Composites for Heat
Exchangers
• Validated Hypersonic
Combustion Codes
Ramjet / Scramjet Engines
Execution : NASP JPO, Contractor, GWPs
Funding : PE 83269F end NASA
T_-'_ P,or',1 .2 . 94 "I_ ,7 _ 090o oI Io::,rTul
x-,lo
I('_) ('41 |3) I*)
s,'m*_smx:_m'
ktalm_,, mr
sxsl
Fl_ Tm_
• Phase $ Funding Eslimete Provided by Air Stiff
• Actual Program Funding Requirement Due 2ncl QTR FY92
I. Concept Selection (4/91)
2. Slu Fresm (2/92)
3. Technology Freeze Date (1/94)
4. Engine Delivery (4/97)
S. Material Flight Engine #1 Selection (1/93)
6. Structure Component Test (3/93)
22-6
ORIGINAL PAGE IS
OF POOR QUALITY
• Develop Advanced Platolet
Rocket Thruster
• Fully Reualblo, Throffiaeble
• High Performance 2-D Rocket
Demonstrates ASO SEC ISP
• NASP Modular Platelet Engine
Selected for SDIO SSTO Concept
• Reliable Electric Restart Via
Lleer ignition System
Advanced Auxillary Propulsion
Execution : NASP JPO, Contractor, GWPs
Funding : PE 63269F and NASA
FurldJll o($M)" :lip _J i2 i._e tit tel !o4 20| z?J IT.2,27.1 lye lOd Ill
_'__ PriorIb9192 93:94 9S 96 97 98 99 O0 01 02 TOtal
In.imPort
FhQhtTest i(1) _X3',(4)
Vlmclal.3n L
(s)
* Phaze 3 Funding Estimate Provided by Air Staff
* Actual Program Funding Requirement Due 2nd QTR FY92
Milestones:
1. System Design Requirements (94L1)
2. Rocket Configuration Freeze (9412)
3. System Preliminary Design (4413)
4. Technology Freeze Date (14)4)
5. X-30 First Flight (10/97)
HIGH SPEED AIRBREATHING
PROPULSION SYSTEM
Alrr BODY
NOZZLE
EXPANSION
SRAMJET ENGINE
FLIGHT MACH 6-ORBIT
:OMPRESSION
SYSTEM HIGH SPEED ENGINES
RAMJET ENGINE
FLIGHT MACH 2-6
FUEL
22-7
O_|SfNAL PAGE iS
OF POOR QUALITY
PROPULSION MODE COMPARISON
5OO
ALT FT
XlO00
0
8 16 24 32
NR
MACH NUMBER OXYGEN
SHUTTLE/ROCKET
NASP
22-8
ORIGINAL PAGE IS
OF POOR QUALITY
NASP X-3O Propulsion Technology Status (Industry)
D. Kenison
NASP JPO
(Paper Not Received in Time for Printing)
23-1



NASP X-30 Propulsion technology status

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