Propulsion challenges and opportunities for high-speed transport aircraft

PROPULSION CHALLENGES AND OPPORTUNITIES FOR HIGH-SPEED TRANSPORT AIRCRAFT W i l l i a m C. Strack ABSTRACT For several years there has been a growing interest in the subject of efficient w sustained supersonic cruise technology applied to a high-speed transport aircraft.

I This presentation identifies the major challenges confronting the propulsion com- co 0-l h munity for supersonic transport (SST) applications. Both past progress and future m I opportunities are discussed in relation to perceived technology shortfalls for an w economically successful SST that satisfies environmental constraints.

A very large improvement in propulsion system efficiency is needed both at supersonic cruise and subsonic cruise conditions. Toward that end, several advanced engine con- cepts are being considered that, together with advanced discipline and component technologies, promise at least 40-percent better efficiency than the Concorde engine.

The quest for higher productivity through higher speed is also thwarted by the at the higher lack of a conventional, low-priced fuel that is thermally stable Extending Jet A-type fuel to higher temperatures associated with faster flight.

temperatures and the adoption of liquefied natural gas (LNG) or methane are two possibilities requiring further investigation.

Airport noise remains a tough challenge because previously researched concepts fall short of achieving FAR 36 Stage I11 noise levels. Innovative solutions may be necessary to reach acceptably low noise.

While the technical challenges are indeed formidable, it is reasonable to assume that the current shortfalls in fuel economy and noise can be overcome through an aggressive propulsion research program.

6-33 CHALLENGES TO HIGH-SPEED TRANSPORTS Although the Concorde ushered in the supersonic transport era, it has n o t been a commercial success for a variety of reasons. Its poor fuel consumption ( 3 times equivalent technology subsonic airplanes) is largely responsible for its uncompet- itive economics; the total operating cost (TOC) is twice that of similar technology, long-range subsonic transports. Very large airframe and propulsive efficiency improvements will be required to alter this situation. In our quest for greater productivity through increased speed, we are confronted with an ever increasing technical challenge arising from high ram temperature levels. In addition to airframe skin temperature problems, the inability of readily available, low-cost fuels to provide adequate thermal stability seriously impedes the pursuit of higher speeds. Expensive JP-type fuels reach thermal stability limits at approximately Mach 3-1/2, but low-cost Jet A is limited to only Mach 2+. While both sonic boom and airport noise levels are currently excessive, only the airport noise problem is of primary concern to the propulsion industry. Another potential environmental issue is the depletion of atmospheric ozone via jet engine exhaust-gas emissions.

CHALLENGES TO HIGH-SPEED TRANSPORTS

ECONOMICS HIGH TEMPERATURES -THERMAL JP FUEL CAPABILITY( STABILITY I I I I 0 1 2 3 4 5 FLIGHT MACH NUMBER -...

OZONE DEPLETION UNCERTAINTY \ NOISE SONIC BOOM PREVENTS OVERLAND FLIGHT v AIRPORT NOISE

I

CD-87-20819 6- 34 SST PROPULSION PROGRESS Considerable progress was achieved during the 1970's in the NASA-sponsored variable- Compared to the 1971 GE4 afterburning turbojet (ABTJ), cycle engine (VCE) program.

the 1981 VCE's consumed 10 percent less fuel at supersonic and transonic conditions, and 25 percent less at subsonic speeds -- reflecting the cycle-changing feature of Nevertheless, VCE's. A simultaneous 25 percent reduction in engine weight occurred.

these gains are insufficient by themselves t o enable competition with subsonic air- craft. The subsonic efficiency of the 1981 VCE engines, for example, is still only one half that of today's high bypass-ratio turbofans.

SST PROPULSION PROGRESS

MACH 2.6 SUPERSONIC

1 .o

1971 GE4 AFTERBURNING TURBOJET RELATIVE 'RANSONIC .9

- ---.. SUBSONIC

RELATIVE FUEL CONSUMP- .8 TlON 1981 VARIABLE CYCLE ENGINE .7 .75 .6 CD-87-28820 .5 6-35 FUTURE HIGH-SPEED PROPULSION PERFORMANCE POTENTIAL The primary cause of the Concorde's high fuel consumption is the dramatic fall in airplane lift-to-drag ratio (L/D) at supersonic speeds which is on the order of 1 / 2 that of subsonic transports. This is only partially offset by the trend toward increasing overall engine efficiency with flight speed. "Installed cruise effi- ?

ciency" shown here includes inlet and nozzle losses, but not nacelle drag, and repre- sents design point values. The middle curve indicates that significant improvement ossible with today's available technology for both subsonic (maximum efficiency - is E 5 technology) and supersonic regimes.

The top band projects future opportunities based principally on NASA cycle analyses.

Several alternative cycle concepts are represented, including very advanced VCE and turbine bypass engines (lower boundary), and radically different concepts such as regenerative air turboramjets (ATR's) and supersonic throughflow (SSTF) turbofans (upper boundary). These advanced technology concepts extend the peak propulsion- efficiency levels from Mach 2+ to at least Mach 4. Gains of 40 percent or more over Concorde's Olympus are possible. Using a simple criterion such as design point efficiency is insufficient to properly convey overall impact. For example, this plot shows a relatively modest 8-percent gain between 1987 technology VCE's and advanced VCE's (lower line of top band). Not shown, but also important are even larger gains in climb efficiency and weight for advanced VCE's.

FUTURE HIGH-SPEED PROPULSION PERFORMANCE POTENTIAL

AIR TURBORAMJET SUPERSONIC FAN INSTALLED

y x E - '87 TECHNOLOGY

CRUISE 4 ' '

READINESS EFFICIENCY, 40

i ------------

q, O h M E CURRENT FLEET SUBSONIC TURBOFANS - THRUST X VELOCITY r l = FUEL POWER I I I I I 1 I CD-87-28821 6-36 VARIABLE-CYCLE ENGINE GOAL The most obvious contender for a future SST is an advanced variable-cycle engine.

This approach builds on the previous VCE philosophy of mitigating the off-design compromises inherent in a fixed-geometry engine. This is accomplished by incor- porating enough variable geometry features to yield respectable performance over a wide range of flight speeds and power settings.

Displayed here is an example of a "goal" VCE, representing what payoffs would accrue if revolutionary advances in materials and structures technology are achieved. This particular design was generated by General Electric in their recent NASA-sponsored Revolutionary Opportunities for Materials and Structures (ROMS) study. It assumes essentially uncooled stoichiometric engine materials coupled to advanced aerodynamics and structural design technologies. This implies extensive use of nonmetallics and :intermetallicmaterials.

Two levels of technology are quoted here: (1) the full stoichiometric goal level is denoted by the right-hand values (GE ROMS), and ( 2 ) a 600 OF cooler level is denoted by the left-hand values (NASA estimate). One-third of the 28-percent fuel reduction is due to a 45-percent engine weight reduction relative to a hypothetical 1984 technology-readiness baseline engine.

VARIABLE-CYCLE ENGINE GOAL

POTENTIAL MACH 3 CRUISE CONDITIONS

"F \ \ / ' T i = 2800 TO 3400 FAN I "F \ \ \ / DRUM / BPR 2 TO 2.4 COMPRESSOR OPR 9 TO 16 UNCOOLED BLADES

BENEFITS (MATERIALS AND AERO) : 290 PAX 5000 nmi TRANSPORT

RELATIVE TO CURRENT TECHNOLOGY AT $l.OO/gal.

[FUEL 24 TO 28%) DOC 17 TO 20% CD-07-20022 6-37 SUPERSONIC THROUGHFLOW FAN ENGINE One potential SST breakthrough is the supersonic fan concept. Instead of using a long and heavy inlet system to efficiently decelerate the intake airflow to the subsonic speeds required by conventional turbomachinery, the supersonic fan effi- ciently processes air at supersonic throughflow velocities. The advantages include much lower inlet-system weight, lighter fan (less stages required for a given pres- sure ratio), less boundary-layer bleed drag, better inlet pressure recovery, and Of course, there better matching of bypass ratio variations to flight speed (Mol.

are many unknowns and challenges. What are such a fan's low-speed operating charac- teristics? How can the core inlet losses associated with unsteady, swirling, super- sonic inflow be controlled; or is an aft fan configuration a better solution? Little effort has been expended on this concept to date, although NASA has initiated a con- cept feasibility research effort.

;- SINGLE-STAGE SUPERSONIC FAN

SUPERSONIC ,,-- 3-STAGE CONVENTIONAL FAN I I r- , DUCT NOZZLE

DIFFUSER ---..

SUBSONIC SUPERSONI DIFFUSER CONVENTIONAL TURBOFAN SUPERSONIC THROUGHFLOW FAN

SUPERSONIC THROUGHFLOW FAN

ENGINE FEATURES I M PLICATIONS

SHORT, ALL SUPERSONIC INLET

LOWER WEIGHT, LOWER INLET DRAG

SINGLE-STAGE SUPERSONIC FAN LOWER WEIGHT AND COST,

RUGGED BLADING

BPR DECREASES WITH Mo HIGHER CRUISE THRUST

CD-87-20023 6-38 BENEFIT OF SUPERSONIC THROUGHFLOW FAN The potential payoff of supersonic throughflow fan (SSTF; technology for a typical SST application has been analyzed by NASA in-house (NASA TM-100114). One of the major contributors is the inlet size and weight reduction to about 1/2 that of a conventional supersonic inlet. This also reduces the inlet bleed-drag penalty.

Furthermore, the higher SSTF inlet recovery leads to more thrust/airflow at cruise, and less transonic-spillage drag when external compression inlets are used. The 35-percent larger cruise thrust/airflow could mean a smaller engine is required dependent on the engine-sizing criteria. In the payoffs quoted here, takeoff thrust/ weight was held fixed to maintain good takeoff performance.

BENEFIT OF SUPERSONIC THROUGHFLOW FAN

MACH 3 COMMERCIAL TRANSPORT

300 PASSENGERS, 5500 nmi RANGE

r_._-.

\ PROPULSION i \\

\ SYSTEM /

WEIGHT i ’/ PERCENT SUPERSONIC FAN IMPROVEMENT AIRPLANE

\ FUEL SAVED

BASELlN E: NON-AFTERBURNING TURBOFAN

- I - I

\ 1 CD-07-20024 6- 39 ADVANCED SUPERSONIC TRANSPORTS COULD ACHIEVE COMPETITIVE FUEL ECONOMY This chart displays the impact of potential future technology advances on airplane fuel consumption while recognizing that the key to viable SST economics is fuel cost levels approaching those for future subsonic airplanes. Achieving 100-percent fuel- usage parity with the subsonic competition is not necessary because of the increased .

productivity associated with SST's. However, it is important to at least be in the same neighborhood, which the Concorde and previous SST-study airplanes cannot achieve despite their shorter ranges. The impact of advanced propulsion technology is * impressive, enabling fuel-consumption rates not much different than current long- range subsonic airplanes. Coupling the most optimistic propulsion technology with potential airframe advances in L/D and structural weight (Wstr) produces encouraging results in the Mach 2 to 4 range. Of course, these are preliminary, first-order results subject to refinement as the ongoing studies evolve. Another uncertainty is the possible introduction of a very advanced, all-new subsonic airplane. An estimate of that possibility is included here that has an 11-percent L/D improvement, a 15-percent structural weight improvement, and a 33-percent propulsion-efficiency improvement. The conclusion to be drawn from this analysis is that the SST fuel- consumption impediment can be overcome, but it will require very large technology gains in all disciplines -- propulsion, aerodynamics, and structures.

IMPACT OF TECHNOLOGY ON FUEL ECONOMY

300 PASSENGERS

+ ADVANCED PROPULSION: SSTF FAN

1 1971 SST h

5500 nmi FUEL Ib 3500 nmi //

SEAT-nmr

ADV. AIRFRAME

+ 20% U D -25%

Wstr 6500 nmi

.2 t

SUBSONIC

I AIRCRAFT

.l t

0 NEW INCREASED PRODUCTIVITY I 1 I 1 2 3 4 5 CRUISE MACH NUMBER CD-87-28825 6-40 SUPERSONIC INLET PERFORMANCE Commercial supersonic flight at Concorde speeds (Mach 2 ) can be viewed as relatively straightforward and within industry's technological grasp. Pushing the cruise speed substantially higher is certainly desirable, but introduces a series of ever- increasing technological challenges. One of these new challenges is illustrated here. Conventional external compression inlets accomplish all of their diffusion outside of the intake duct through several oblique shocks and a terminal normal shock located at the cowl lip. This type of inlet delivers adequate performance and is well-behaved (stable) under all transport flight conditions up to Mach 2 . Beyond Mach 2 though, the performance of external compression inlets rapidly deteriorates because of the excessive cowl drag associated with the increasing cowl-lip angle and the inability to increase the number of oblique shocks because of excessive inlet length and weight penalties. Flight beyond Mach 2, therefore, requires a mixed- compression-type inlet that performs some of the diffusion inside the intake duct through more oblique shocks and a normal shock near the throat. This introduces other problems: notably, more boundary-layer bleed to avoid adverse shock-boundary- The result is a layer interactions (separation) and inlet shock-system instability.

much more complex inlet and control system. Neither transports nor fighters have been flown operationally with such inlets, yet the need for utmost propulsion effi- ciency will require it for high-speed transports.

SUPERSONIC INLET PERFORMANCE

1 .o

TOTAL PRESSURE .6 RECOVERY .4 COMPLEX:

- STABILITY (UNSTART)

LOW PERFORMANCE ABOVE MACH 2

- CONTROLSNARIABLE GEOMETRY

- PRESSURE RECOVERY

HIGHER WEIGHT

- COWL DRAG

HIGHER BOUNDARY-LAYER BLEED DRAG .2 0 1 2 3 4 5 6 DESIGN MACH NUMBER CD-87-28826 6-41 MIXED-COMPRESSION SUPERSONIC INLET INSTABILITY Mixed compression inlets are quite susceptible to a phenomenon kncwn as inlet instability o r "unstart." Whenever a flow-retarding disturbance occurs, the internal shock system moves abruptly upstream and repositions itself completely outside the intake duct. This causes an abrupt and severe drop in thrust due to lower recovery and mass flow, and an increase in drag. The precipitating disturbance could be relatively small, such as encountering a strong gust or rapidly changing the angle-of-attack. If the initial disturbance is large (e.g., compressor stall),

the transient response can be very severe -- possibly unstarting neighboring inlet-

engine systems which would likely throw the airplane into a violent yaw and roll maneuver. To prevent such undesirable behavior, some form of stability control system is needed.

MIXED COMPRESSION SUPERSON IC IN LET I NSTABl LlTY

J STABILITY MARGIN TRANSIENT BOW WAVE .9 DESIGN POINT

/w

/' UNSTART PATH,,,' & . .

.8 CROSS UNSTART SEVERE GUST OR Aa ENGINE TRANSIENTS

t

- COMPRESSOR STALL

PRESSURE

- AUGMENTOR IGNlTlOl

RECOVERY OR BLOWOUT .7 INLET UNSTARTE .6 ABRUPT THRUST LOSS I, ABRUPT DRAG INCREASE INLET BUZZ I I I I I

.6 .7 .8 .9 1 .o

MASS FLOW RATIO 0-87-28827 6-42 MIXED-COMPRESSION INLET STABILITY IMPROVEMENTS This inlet stability improvement concept consists of a set of self-actuating bleed valves located in the inlet nacelle. These rapid-response-rate pneumatic valves will open in response to the increase in duct pressure produced by a transient excursion of the inlet terminal shock forward from its steady-state position. A s the shock moves forward it exposes the stability bypass plenum to increased pressure and auto- matically activates the bleed valves which spill inlet bleed air overboard. This increases the inlet mass flow and forces the shock rearward, and thereby reestab- lishes stability. The valves close when the transient disturbance subsides and the shock has retreated to its steady-state position.

An experimental wind tunnel test program at NASA Lewis Research Center verified the feasibility of this concept during the mid 1 9 7 0 ' s . A five-fold increase in stability margin was demonstrated using a YF-12 system simulation.

Considerable research lies ahead, however, to adequately address this important issue.

MIXED COMPRESSION INLET INSTABILITY IMPROVEMENT

1.0 INCREASED MARGIN WITH STAB. AUGMENTATION BASELINE

1 . 1 1 STABILITY

.9 DESIGN POINT FAST-ACTING VALVES,7 COWL BLEED REGION-, .8

e

AIRFLOW , , PRESSURE RECOVERY .7 WIND TUNNEL TESTED .6 5-FOLD INCREASE IN STABILITY MARGIN .5 I I I I 1

.6 .7 .8 .9 1 .o

MASS FLOW RATIO CD-07-20020 6-43 NOZZLE PERFORMANCE The exhaust nozzle for an SST must perform well at three critical flight conditions -- takeoff, subsonic cruise, and supersonic cruise. These experimental model test results (Lewis Research Center, 8- by 6-ft wind tunnel) of an ejector nozzle show that, while good takeoff and cruise performance was achieved, the subsonic cruise performance was disappointingly low because of flow separation over the inlet doors of the ejector. This shortfall is important because it significantly increases the reserve fuel allowance required to reach an alternate airport -- and, for long-range SST's, the amount of reserve fuel is quite large. In addition, it is critical to obtain high nozzle performance at the transonic thrust minus drag "pinch point" to minimize inlet-engine flow matching penalties.

NOZZLE PERFORMANCE

VCE RESEARCH EJECTOR

1 .oo

.98 .96

- ---e--- SCR GOALS

GROSS .94 THRUST COEFFICIENT, -92 - Cf .- .90 - SUBSONIC .88 CRUISE A .86

0 .5 1 .o 1.5 2.7

MACH NUMBER 6-44 TRANSONIC PROPULSION SYSTEM DRAG Just as exhaust nozzle performance is critical during transonic flight, so also is the minimization of transonic installation losses associated with inlets and nozzles.

The transonic inlet spillage drag, for example, can exceed the entire airframe drag €or high design Mach numbers. This problem arises from a major mismatch in inlet €low-swallowing capacity (too much) compared to the engine demand. Likewise, the nozzle boattail drag penalty rises rapidly with design cruise speed. Finding solu- tions to these installation problems is absolutely essential to achieve an acceptable airplane design.

TRANSONIC PROPULSION SYSTEM DRAG

INLET NOZZLE

SUPERSONIC CRUISE TRANSONIC SPILLAGE

1.0 C

-'t

0 1 I I 2 3 4 0 1 2 3 4 FLIGHT MACH NUMBER CRUISE MACH NUMBER 6-45 THE HIGH-SPEED TRANSPORT FUEL ISSUE Conventional jet fuels cannot withstand the high temperatures associated with flight speeds in excess of about Mach 2 . If subjected to temperatures above approximately 250 OC (time dependent also), they thermally decompose and form coke deposits that clog fuel-supply components. Consequently, a challenge exists to extend the thermal stability of conventional jet fuel (Jet A ) to higher temperatures without incurring a significant fuel price increase -- either in the fuel manufacture o r associated with special fuel transportation and handling requirements (such as with JP-7 and c cryogenics). While the practical use of hydrogen lies far into the future, liquid methane or LNG remains as an intriguing possibility because of its current low price and high thermal stability. Endothermic fuels offer more heat sink capacity, but are fraught with offsetting practical and economic penalties. Uncertain future fuel prices and infrastructure costs cloud the issue of fuel selection and, consequently, airplane design speed as well.

THE HIGH-SPEED TRANSPORT FUELS ISSUE

CONVENTIONAL HYDROCARBONS RELATIVE FUEL PRICE STABILITY LIMIT

1 I I I I I 1

0 1 2 3 4 5 6 MACH NUMBER CD-87-28831 6-46 PROGRESS IN SST TAKEOFF-NOISE REDUCTION The first generation of hypothetical U.S. SST’s of the early 1 9 7 0 ’ s used after- burning turbojets and would have provoked the irritation of many people living around major airports. Reducing their high jet exhaust velocities (over 4000 ft/s) by oversizing the engines and throttling back during takeoff reduces noise somewhat, but it also increases airplane size too rapidly to be an effective method for more than a few dB. Each curve represents a series of various amounts of engine over- sizing for a fixed mission. Considerable noise reduction progress evolved during the 1970‘s through a combination of variable-cycle features and many noise suppression concepts experimentally tested. However, even this progress is insufficient to meet current FAR 36 Stage I11 requirements. Much research lies ahead if we are to achieve a quiet SST without excessive noise reduction penalties.

PROGRESS IN SST TAKEOFF NOISE REDUCTION

MACH 2.4 TO 3.2 EXPERIMENTAL DATA BASE

STAGE 3 EQUIVALENT STAGE 2 I

I \

I I I 1971 SST I I I GE4 AFTERBURNING I I I

1987 \ I \ TURBOJET

I I I i VCE WITHOUT I

I SUPPRESSION

I TAKEOFF

\ \

ENGINE I GROSS EJECTOR CoAN LAR I OVERSIZING I

SUPPRESSOR AND CHUTE NozzK \

WEIGHT, I I k Ib 750

I \

I I I I I , I I I 1

700 I 1

100 105 110 115 120 125 SIDELINE EFFECTIVE PERCEIVED NOISE AT 1500 FT, dB

CONCLUSION: CONSIDERABLE RESEARCH EFFORT WARRANTED

CD-87-28832 6-47 J E T NOISE REDUCTION CONCEPTS Some of the noise reduction concepts illustrated here have been explored in axisym- metric configurations suitable for Mach 2-3 airplanes. These concepts need data base extensions for two-dimensional nozzles suitable for higher flight speeds. Other con- cepts have practically no data base at all and are quite speculative. For example, the concept of cancelling source noise by superimposing an out of phase second source has made significant strides recently and appears suitable for discreet frequency noises such as produced by a propeller. Extending this idea to cancel broadband jet noise with passive secondary noise sources (pneumatic oscillators) represents a very speculative and technically challenging strategy. The remote augmented thrust system concept guarantees low noise with its high mass flow, low pressure ratio fan. But it introduces different problems -- notably, how to integrate the remote deployable takeoff fans into the airframe.

JET NOISE REDUCTION CONCEPTS

1 rOUTER STREAM

ACOUSTI -L

I - -

/ \ SUPPRESSOR INVERTED VELOCITY PROFILE ACOUSTIC LINING PRIMARY- %/, SHIELD STREAM

SUBSONIC CRUISEITAKEOFF J (HIGH TEMPERATURE

OBSERVER Low VELOCITY) / \ d THERMAL ACOUSTIC SHIELD REMOTE AUGMENTED EJECTOR THRUST SYSTEM CD-07-20033 6-48 PROGRESS IN SST CRUISE NOx-EMISSION REDUCTION Presently, it does not appear that we have a known problem with SST engine emissions.

There is some concern, however, that we might have a future problem if ongoing analyses conclude that significant upper atmospheric ozone depletion would be caused by a fleet of NOX-emitting SST's. Previous airport pollution concerns precipitated a NASA emissions-reduction research program that led to the development of several control mechanisms including two-zone combustors. The 1970's engines had single-zone combustors that had their high-power efficiency compromised to obtain good low-power ignition and stability. The improved two-zone combustors used a pilot stage opti- mized for idle conditions and a main stage optimized for cruise power. This resulted in leaner, well-mixed cruise combustion with approximately one-half as much cruise NOX emission assuming the engine cycle remains unchanged.

However, our continued quest for higher overall engine efficiency produces ever higher cycle temperatures which increases NOX production. Hence, the final engine designs of the supersonic cruise research (SCR)/VCE program, if built, would have produced about as much NOX as the actual engines introduced a decade earlier. Today, we face the same dilemma: performance-driven designs will increase NOx, while emissions-driven designs will reduce performance.

PROGRESS IN SST CRUISE NOX EMISSION REDUCTION

IMPROVED CYCLE (HIGHER T AND P) SINGLE-ZONE COMBUSTOR INVARIANT EMISSION TECHNOLOGY PERFORMANCE DRIVEN WHERE TO?

NOX EMISSION INDEX, g/kg FUEL EMISSIONS DRIVEN IMPROVED EMISSION TECHNOLOGY INVARIANT CYCLE

I 1 1

1970 1980 1990 YEAR CD-87-28834 6-49 NOx-EMISSIONS-REDUCTION CONCEPTS One approach to reduce NOX emissions is to reduce the flame temperature. Another approach is to reduce the residence time of the combustion gas at high temperatures.

In the latter approach, two concepts worth pursuing are (1) increasing the velocity through the combustor, and ( 2 ) avoiding large recirculation regions within the primary combustion zone. Increasing the combustion velocity to relatively high- subsonic values involves finding means to avoid excessive pressure losses, as well as maintaining good combustion stability and ignition characteristics. Avoiding large pockets of recirculating hot gases in the primary zone also reduces stability char- acteristics and, thereby, requires the implementation of other stability-enhancing features.

NOx-EMISSIONS-REDUCTION CONCEPTS REDUCE RESIDENCE

TIME AT HIGH TEMPERATURE

HIGH VELOCITIES I N PRIMARY ZONE I / I PRIMARY PREMIX r SECONDARY FLAMEHOLDER SUPERSONIC

AIR - y COMBUSTION

----,I \ I I FUEL NOZZLES’ iSECONDARY NOx7 9 PREMIX PASSAGE FUEL, kg I

/- I I I

2 4 6 RESIDENCE TIME, ms CD-87-28835 6-50 CANDIDATE HIGH-SPEED PROPULSION PROGRAM PLAN As the 21st century approaches, it is becoming increasingly clear that efficient supersonic cruise flight is within our technological reach. Many challenging propulsion problems need to be addressed, however, before a state of technology readiness is achieved. One possible program plan entails a two-pronged approach: a near-term effort aimed at variable-cycle engine concepts incorporating very aggres- sive discipline and component technologies, and a far-term effort focused on vali- dating supersonic throughflow technology which offers even higher potential benefits.

Continued propulsion system studies as well as a high-speed fuel and fuel systems effort are also needed. Attainment of the propulsion goals outlined herein would indeed revolutionize aircraft capability for the future.

CANDIDATE HIGH-SPEED PROPULSION PROGRAM PLAN

PROPULSION SYSTEM STUDIES 1

I f ? S C I I

HIGH MACH FUEL STUDlESlRESEARCH 1

I

I t LOW-EMISSION COMBUSTORS VARIABLE-CYCLE ENGINE (VCE) TECHNOLOGY VARIABLECYCLE READINESS PROPULSION SYSTEMS COMPONENT INTERACTIONS

I

SUPERSONIC THROUGHFLOW (SSTF) COMPRESSION SYSTEM TECHNOLOGY READINESS

r I

CONCEPT VALIDATION EXPERIMENTS SUPERSONIC

G

THROUGHFLOW TECHNOLOGY CD-07-20036 6-5 1