Uniquely advantageous features, on both the performance and weight sides of the ledger, can be achieved through synergistic design integration of airbreathing and rocket technologies in the development of advanced orbital space transport propulsion systems of the combined cycle type. In the context of well understood advanced airbreathing and liquid rocket propulsion principles and practices, this precept of synergism is advanced mainly through six rather specific examples. These range from the detailed component level to the overall vehicle system level as follows: using jet compression; achieving a high area ratio rocket nozzle; ameliorating gas generator cycle rocket system deficiencies; using the in-duct special rocket thrust chamber assembly as the principal scramjet fuel injection operation; using the unstowed, covered fan as a duct closure for effecting high area ratio rocket mode operation; and creating a unique airbreathing rocket system via the onboard, cryogenic hydrogen induced air liquefaction process

Leingang, J. L.

English

NASA Technical Reports Server

N92-21520
ADVANCED RAMJET CONCEPTS PROGRAM
J. L Leingang
Wright Research and Development Center
Wright Patterson AFB, OH
Uniquely advantageous features, on both the performance and weight sides of the ledger, can be
achieved through synergistic design integration of airbreathing and rocket technologies in the
development of advanced orbital space transport propulsion systems of the combined cycle type.
In the context of well understood advanced airbreathing and liquid rocket propulsion principles
and practices, this precept of synergism is advanced mainly through six rather specific examples.
These range from the detailed component level to the overall vehicle system level as follows:
Utilizing jet compression, as a specific air-augmented rocket mode approach
Achieving a high area-ratio rocket nozzle through innovative use of air-handling
ducting
Ameliorating gas-generator cycle rocket system deficiencies while meeting, (1)
ejector mode afterburner and, (2) rocket-mode internal aerodynamic nozzle
operating needs
Using the in-duct special rocket thrust chamber assembly as the principal scramjet
fuel injection station
Using the unstowed, covered fan as a duct closure for effecting high area-ratio
rocket-mode operation
Creating a unique airbreathing rocket system via the onboard, cryogenic hydrogen-
induced air liquefaction process.
ILLUSTRATIVE CASES-IN-POINT: AIRBREATHING/ROCKET SYNERGISM
.IET COMPRESSION
Jet compressors resemble conventional ejectors as used in industrial applications of yesterday
(steam-locomotlve smokestacks), and today (steam-cycle electric powerplant vacuum
condensers). A propulsion-oriented application familiar to rocket test personnel are steam
ejector systems used to initially "pull down" and/or actively exhaust a rocket altitude-simulation
facility (e.g., as used in RL-10 engine testing).
In the combined cycle engine context, jet compressors are made up of a supersonic primary flow
unit installed in a duct with an inlet providing the secondary flow to be compressed (air in this
case). The downstream portion of the duct is divided into a mixer (usually of constant area),
4-1
https://ntrs.nasa.gov/search.jsp?R=19920012277 2020-03-17T13:11:07+00:00Z
followed by a diffuser having diverging geometry for diffusing the mixed high-subsonic flow.
Compression is achieved in the mixed stream by virtue of the h/gh "driving" enthalpy of the
primary flowstream, in this case a rocket.
Such jet compressors are characterized as "effective," if not necessarily "efficient" compressors,
having characteristic advantages and disadvantages. They are lightweight, rugged, and highly
tolerant of flow-distortion profiles, on the one hand. On the other, in contrast to conventional
turbomachines, they have relatively high propellant consumption rates, and require considerable
mixing duct lengths, hardware which usually must be actively cooled.
AIRBREATHING-MODE DUC'q3NG USED AS HIGH AREA-RATIO ROCKET NOZZLE
The obvious technical approach is to utilize pan of the a/r-handlin 8 duct and the airbreathing
mode(s) combustor/nozzle assemblies to this end. Although this approach remains yet
undemonstrated in "the problem has been solved" sense, there is analytical and even
experimental evidence that this is, indeed, a feasible design approach. This evidence will be
summarized below.
Specifically, the objective is to so configure and operate the engine (in rocket mode) as tO
provide an "aerodynamic or virtual" nozzle extension for the physical rocket unit with its low
expansion-ratio nozzle. This aerodynamic extension would "control" the underexpanded
supersonic exhaust plume such that it would smoothly attach to the engine's specially configured
airhandling duct now serving as a physical nozzle extension. The objective is to minlmiT¢ shock
losses and otherwise non-optimum intermediate exhaust expansion processes.
Once attached, the divergent final section of the duct would continue the nozzle expansion
process to very high area ratios of, say, several hundreds-to-one. The following flow exiting from
the engine duct, or nozzle, further expansion of the rocket exhaust would take place on the
vehicle aft-body. Nozzle aerodynamics-wise, the mechanization of th/s latter approach is seen to
be a fortunate carry-over from the supersonic combustion ramjet mode (where aft-body
expansion is a virtual necessity), assuming that the scramjet mode is to be used.
ADVANTAGEOUS DISPOSITION OF ROCKET SUBSYSTEM
TURBOPUMP GAS GENERATOR EXHAUST
CONVENTIONAL LIMITATIONS OF GAS-GENERATOR TYPE ROCKET ENGINES
Historically, hydrogen/oxygen rocket engines have utilized turbopump propellant feed delivery
systems. This achieves high combustion pressures leading to advantageous high area-ratio nozzle
operation, without the structural weight penalties which might accompany pressure-fed systems,
which are inherently difficult to execute with liquid hydrogen fuel
Various turbopump drive approaches have been selected in hydrogen-oxygen rocket engines
developed to date. For U.S. engines developed so far, the following turbopump drive
approaches have been used:
4-2
RL-IO
J-2
SSME
Expander Cycle
Gas Generator Cycle
Staged Combustion Clopping) Cycle
Looking ahead, it would seem to be the case that the staged combustion system will continue to
be favored for large rocket engines, e.g., possible successors to the SSME (which pioneered this
turbopump drive cycle). For smaller engines, such as those which might be applied to orbital
transfer vehicle (OTV) systems, the expander cycle appears to be favored. Why not the gas
generator cycle for these applications? The answer is, its lower special impulse performance, as
nest discusses.
Although the gas generator cycle has a number of technical advantages (e.g., low pump-out
pressures, achievement of high engine thrust/weight ratios, reduced interfacing difficulties), it
has one salient and intrinsic disadvantage: a low-pressure fuel-rich turbopump turbine exhaust,
which is not very effective in producing additive thrust. This leads to an overall engine specific
impulse deficiency in comparison with the competing turbopump-drive approaches.
GAS-GENERATOR CYCLE IN THE COMBINED CYCLE ENGINE
In addition to its advantages in the conventional rocket engine context, which should carry over
into a combined cycle system, the gas generator cycle may actually be strongly preferred in
selecting the design of the engine's rocket subsystem. The main reason ties in unique uses of
the turbopump-drive fuel-rich exhaust flow. A secondary reason for this preference lies in what
is probably a poor physical design integration prospect for both the staged-combustion, and
expander cycle alternative.
ROCKET MODE
Taking the second-name mode first, it has been previously treated as the probable need for a
finite secondary flow (Rocketdyne's "basebleed") to maximize "aerobell" nozzle performance.
The turbopump exhaust of the gas generator cycle seems to be a natural source of the base-
bleed flow. In fact, the secondary flow might well have to be otherwise created in the case of
the staged combustion cycle, for instance.Whether the gas generator exhaust flow "naturally
matches up" with the special aerobell configuration needs in its quantity available, or in its flow
properties, has yet to be examined, perhaps some system optimization effects will be devised to
steer the mrbopump component design in one direction, or another. But the point remains that
there is what appears to be a good"fit" for the gas generator cycle (and not the other
alternatives) in mechanizing the combined cycle rocket mode. After all, this is the implied
expectation supported by the effort reported by the Rocketdyne researchers.
EJECTOR MODE
Proceeding now to the first of the two rocket-operating modes, the ejector mode, some
discussion of the thermodynamic process nature of this operation is called for. This has to do
with the specific type of air-augmented rocket to be selected, as covered in the earlier discussion
of the jet compressor process.
4.3 ¸
Alteraative concepts proposed for accomplishing air augmentation of rockets, presumably to
raise specific impulse and/or thrust levels, are several. They range from a_fixing a simple,
lightweight fixed-geometry duct or shroud around an otherwise conventional rocket engine, to
more complex, but are judged to be more workable systems. Unforv.mately, space here does not
permit even a summary review of the possibilities. Alternatively, let usgo to one leading-
candidate type system as evidenced in previous assessments the afterbuming cycle air augmented
rocket.
DIFFUSION AND AFIERBURNING CYCLE AIR AUGMENTED ROCKEr
The stipulation of a stoichiometr/c rocket (departure from the usually fuel-rich setting) now joins
the earlier specification of an unconventional "distributed" high shear-area configuration, as a
design precept, or hardware determinant. This is, of course predicated on the selection of the
DAB cycle.
Now the afterburning aspect of this cycle implies making fuel available in the afterburner
combustor, downstream of the ejector's mixer and diffuser. Typical engine designs provide for
conventional afterburner (or ramburner) fuel supply and injection means. Here is how the gas
generator cycle again appears to fit in well. Recall its performance detriment, as a rocket,
stemmed from its characteristic fuel-rich (to control turbine temperatures) low pressure (turbine-
drive enthalpy extraction) exhaust.
This gas can now be very usefully combusted in the (relatively) low-pressure afterburner.
Calculations check that, even with this hot gaseous fuel supply, substantial amounts of additional
fuel-hydrogen in this iUustration--are required to burn stoichiometr/cally with air, for full-
engine-power conditions, as needed for h/gh-thrust acceleration propulsion operation. Hence,
the gas generator cycle's turbine exhaust "detriment" is largely removed and, as we have seen,
the gas generator cycle becomes uniquely and naturally the system of choice in the type of
combine cycle engine under discussion.
IN-DUCT SPECIAL CONFIGURATION
ROCKET THRUST CHAMBER ASSEMBLY FOR SCRAMJET FUEL INJECTION
A specialized "rocket engine subsystem" was designed by Rocketdyne for the 1966-1967
"Composite Engine Study". The special combined cycle engine for which this subsystem was
designed was a circular cross-section version of what is referred to as the ScramLACE engine.
It turns out that the f'mal version of this engine had a rectangular cross-section duct
configuration to more optimally match up with the Lockheed-designed first-stage vehicle of the
study. The 2-ring configuration was, in the final version of this engine accordingly replaced with
eleven vertically-mounted "linear" thrust chambers of equivalent propellant flow-rate and thrust.
These hydrogen-fueled units used liquid air (LAIR) as oxidizer.
4-4
TRUE AIRBREATHING ROCKET SYSTEMSVIA AIR LIQUEFACTION
Onboard cryogenichydrogen-inducedair liquefaction was fairly heavily dealt with in the original
aerospaceplane R&D of the 1960s in both analysis and design experimental set-ups. Some of
the design consequences of this re_gerative capacity limitation are gone into below. As the
available airbreathing achieves testify, a number of air liquefaction cycles were vigorously
explored in the early 1960s, ranging from "BasicLACE through SuperLACE" to "ACES (Air
Collection and Enrichment System)."
These cartier developments were substantially more extensive than today's propulsion engineers
seem to be aware, not just in terms of dollars and manhours expended, but in depth and
sophistication of the design analyses conducted, and the experiments which were run. All in all,
there was considerable success in the early developmental experiments conducted.
By the late 1960s, when there was a cessation of further research and development work in air
liquefaction systems and subsystems, a substantial level of technology had been documented.
Particularly important, some of the salient design and operating challenges became fairly well-
known, and design solutions accordingly brought forward.
One of the salient objectives of liquefying air in such systems was to provide a means for
operating very lightweight, compact engines, which could gain the "airbreathing advantage" while
maintaining rocket-like qualifies of low weight and compactness. Such was achievable in
principle and technically illustrated in small-scale test rigs encompassing: (1) a liquid air pump
of the rocket engine type, and (2) a high-pressure rocket-type thrust chamber. These equipment
items were much more compact and less massive than the corresponding conventional ambiem-
temperature air compressor and its drive turbine and, lower combustion pressure turbojet-type
combustor or afterburner assemblies.
4-5
|¢/I
,,=J
I
k_
i,
,f,J
1-
I-
w
4000
2000_
1ooo
800
• 600
400
EXPANSION TO FREE,JET
STATIC PRESSURE
qns 0.98
I
ct
200 /
/
lO0
O
RAMJET MODE
0 PRWARY'_ AIR
D EJECTOR {A)
•_ EJECTOR (0)
_) RAMJET (B)RA JET (B+C)
FREF..JET VELOCiTY, h/secxlO'_
EJECTOR SCRAM.JETTEST•E_GINE PERFORMANCE
FIXED GEOMETRY
(.1) _'lx(C) CLO" (TRY. Le C_%]mt-CTIDf, RATIO,
ALl.-J. I_1[llk/-L CO*'MI| SSIOE;
(b)
VARIABLE GEOMETRY
..........__J-t1"_'_ 1,,
°°
IIA',tLI|I',{. t,_,,L ._llClpl
It)
;. , - . _- _..
...."°• ..... •.
u
f
(e)
SPECTRUMOF POTENTIALEJECTORSCRAMJET INLETCONCEPTS
4-6
SEC/PRt SN. OYIIAMIC
SYI_. RUN FLOW RATIO AL_TVOE PRESS.
0 37 3.24 45,,200 ]920
O 25 3.1Z 46,500 1800
- 3? S0,000 1525
37 52.500 1560
22 $9,000 990
SOD
450
" 400
..a
I
__ 3S0
1.1
300
PRI. A/B
PRI. EQUIV. EQVUI.
PRESS. RATIO RATIO
4.94 0.97 0.84
4.96 0.96 0.85
2.70 4.9.5 0.96 0.83
us 4.. o,. o.es
UD 4.9S o.gs o.a2
,NUETdAP'ru,_E',.',_ - "78.S'-.i
Q
I
L_
_9
@
J......._
._t
°I
_ 8
2so _8
um i_.
2oo 1 I h
0.3 0.4 O.S t 6 0.7 D,
INLE T PRESSURE RECOVERY RATIO
/_I{)I'B1 EJECTOR RAMJET ALTITUDE PERFOR/_NCE
MACH _.o0 EJECTOR MODE
PRI. A,'B
SEC/I_RI. _I. EQUIV. EQUIV. Sla. OYII_IC
RUN FLOW RATIO I_ESS. RATIO RATIO ALTITUDE PRESS.
37 1.82 496 0.q4 1.00 42,000 900
28 2.08 504 0.94 0.87 38,500 1060
28 2.65 SOl 0.93" 0.87 33,000 !.370
PREDICTED
- 400
__ 3SO
"_ 3oc
v.a
!
MACH ].Y FLIGHT CONDITIONS I
INLET CAPTURE AREA+ = Sq in, 2 I
0_, % = 2000 _sl
i
t
!
,,,
........
' O_ + tooo
' ( +
I
' I
1 il= _
I,I,I lid _Id
_oo/ $ I I I I } J
o.s o.e o.7 o.e o.9 z.o
INLZTPRESSURER COVERYRATIO
MAlqI_B! EJECTOR RAMJD ENGINE PERFORMANCE
EJECTOR RAMJET OPERATING MODE
ALTITUDE OPERATION
4-7
loo i i j Ii ! , iI i
q. J ' . I _ I
[ _:_ 0 o. I I , -
o o i o_......-._'_ : I
t ! : I , I
• : t } SYMBOL RUN
__ 0 11
° "--- I I I _ .
' 0 36
90 "
50 60 70 80 90 lO0 110 120 130
MOK_ENTUM RATIO, PRIMARY/SECONDARY
MAI91-BI EJECTOR RAMJET ENGINE PERFORMANCE
JET COMPRESSOR PERFORMANCE VERSUS MOMENTUM
RAT I0 - PR IMARY I SECONDARY
Z
U
_J
I--
Z
W
Z
o
U
100
¢
90
d
70
bO
5O
i
40
3O
/'I !s%.oL
! 0
I
I
i
-0
COMPONENT "
PRIMARY
JET COMPRESSOR
DIFFUSER
AFTERBURNER
200- 1 2 3 4 s
MIXER LENGTH/DIAMETER RATIO
MAI91-BI EJECTOR RAMJE'I ENGINE PERFORMANCE
COMPONENT EFFICIENCY VERSUS MIXER LENGTH/DIAMETER RATIO
4-8
10000
3000
oOOO
4000
2DOC
W
1000
800
bOO
400
20C
I ! I I
-- IOEAL PERFORMM,/CE
...... EJECTOR SCRAMJET ENGINE
FIXED iNTERNAL GEOMETRY
VARIABLE INLET
FULLY EXPANDED EXIT NOZZLE
I
I ,
t
f
,/5"-'" I
• EJECTOR/
./
MODE
T/KE : o.g2
SUBSONIC COMBUSTION
._ .: 1.0
W2 /wp- VARIABLE
SCRAMJET MODE
--_E = 0.985I
SUPERSONIC COMBUS1
• VI: o
10 12
FLIGHT MACH NUMBER
COMPARISON OF EJECTOR SCRAM.JET CYCLE PERFORMANCE
CHARACTERI STICS
4-9_



Advanced ramjet concepts program

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server