Carl A. Trexler- Performance of an Inlet for an Integrated Scramjet Concept

8/3/2019 Carl A. Trexler- Performance of an Inlet for an Integrated Scramjet Concept

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SEPTEMBER 1974 ENGINEERING NOTES 589aircraft is dry washed or polished. Total lapse time is ap-proximately six hours. This accomplishment is credited tothe specialized equipment for gaming easy and rapid ac-cess to all areas of the exterior surface.Several operators are evaluating a barrier coating overthe areas of high soil exposure. This coating traps the soilson the surface without permitting them to embed in thepaint. At the washing cycle, the barrier coating is releasedfrom the airplane surface by a chemical solution which isspray applied. With the release of the barrier coating, thesoils are removed, leaving the regular painted surface fair-ly easy to clean by the conventional cleaning materialsand methods. The barrier coating is then reapplied.Another two operators are reported to be flight testing aclear polyurethane topcoating which is applied over glos-sy, pigmented paint. Purpose of the test is to determine ifthe topcoat will prolong gloss retention of the pigmentedsurface and if it will enhance cleaning time and effort .The Military Aircraft Command is planning indoor fa-cilities for complete washing of its C-5A fleet. Outdoor fa-cilities will continue to be used, but hopefully limited tothe washing of lower surfaces which are accessible fromportable stands and standard length brushes and mops.

Fraternal Approach Toward AlleviationProcedural problems are surmountable for effective, ef-ficient cleaning of jumbo jets. In time, special equipmentwill be developed, as will satisfactory cleaning com-pounds; nevertheless, much of the developments will haveto be tailored to specific aircraft configurations, cleaninglocations, schedule requirements, and operational spec-trum. For resolutions to the procedural problems forwashing mammoth aircraft, contributions must be appliedwillingly from the experiences of aircraft manufacturers,operators, washing contractors, and the specialized chemi-cal compounders. Testing and evaluating of special prod-ucts and equipment are of interest to all operators; conse-quently, a mutual organization should have this informa-tion for assessing materials and procedures and for recom-mending improvements. Such an organization is the AirTransport Association, which has an active committee foraircraft cleaning. This committee has both civil and mili-tary participation. By the active participation of its mem-bers, uniformity of procedures will be promoted so thateconomy can be shared by all contributors. By collectiveaction of concerned and technically capable people, theprocedural problem of washing the jumbo jet will be re-solved.

Performance of an Inlet for anIntegrated Scramjet Concept

Carl A. Trexler*NASA Langley Research Center, Hampton, Va.

IntroductionDURING the past decade, the USAF and NASA havefunded the development of several small-scale researchscramjet engines. These projects have shown that the

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C O M P R E S S I O N S U R F A C E - 7 V ~

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E N G I N E M O D U L E S- F U E L I N J E C T I O N STRUT/ - T H R O A T S

Received Apri l 15, 1974.Index categories: Airbreathing Propulsion; Hypersonic.*Aerospace Technologist, Hypersonic Vehicles Division, A d-vanced Concepts Section.

U P S T R E A M E N D O F N O ZZL E- I N L E T LC O M B U S T O R

Fig. 1 Inner module of Langley Scramjet Module Concept.4

scramjet is a feasible engine concept; practical levels ofthrust have been demonstrated and a substantial technol-og y base has been established.1 NASA is now conductinga hypersonic (Research and Testing) program2'3 which isdevoted in part to the next logical step in scramjet evolu-tion, the development of engine concepts which will inte-grate with the airframe. Integration includes the use ofthe vehicle forebody to precompress the engine airflow be-fore it enters the inlet and the use of the vehicle afterbodyfo r additional expansion of the nozzle exhaust gas. Otherprincipal design criteria are low engine cooling require-ments to make part of the heat sink of the hydrogen fuelavailable for active cooling of the a i r f rame, fixed geometryto reduce weight and system complexity, and minimumexternal drag. Detailed design studies utilizing advancedbasic technology have resulted in the definition of aunique engine concept (Fig. 1 and Ref. 4), which conser-vative predictions indicate will meet all the above designobjectives. Innovative design features of the inlet andcombustor coupled with the favorable effects of integra-tion permit high levels of performance over the Machrange from 4 to 10 with relatively low cooling require-ments. For instance, at Mach 6 a specific impulse ofabout 3000 sec is predicted with only 40% of the fuel heatsink required for engine cooling. Experimental investiga-tions are now in progress to substantiate and further de-velop the design concept; the present Note reports brieflyon the measured performance of the inlet at Mach 6.

Inlet Design ConceptThe cross section of the Langley Scramjet module (Fig.1) varies from a nearly square capture area to a rectangu-lar inlet throat to a square combustor exit. This type of

configuration favors low cooling requirements by reducingthe internal wetted area. In addition, a cluster of severalsuch modules mounted on the underside of the vehicle iscapable of capturing all the airflow lying between the ve-hicle surface and the vehicle bow shock at the maximumMach number, thus producing a maximum thrust. Wallfuel injectors would produce very long mixing lengths fo rthis configuration; therefore, three fuel injection strutshave been provided to allow six planes of fuel injection inthe stream. This feature not only shortens the combustorbut also the inlet since the struts provide a significantpart of the inlet flow compression. The leading edges ofthe sidewalls and all downstream stations are swept at 48to provide spillage at low Mach numbers for starting withfixed geometry. Spillage occurs through the open windowupstream of the cowl leading edge, which is bathed byshocks from the sidewall compression surfaces. The com-bination of the sweep angle, the sidewall design, and thecowl leading edge location produces near-maximum masscapture ratios as a function of Mach number, and thespilled flow provides a lift increment for the vehicle. The


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1!Fig. 2 Inlet model and Mach 6 shock diagram.

swept shock system also provides less compression nearthe vehicle imderbody and allows the inlet to ingest thevehicle forebody boundary layer without separation. Theexternal surface of the cowl is parallel to the local bodyline of the vehicle and has a very small drag.The photograph in Fig. 2 shows one inlet sidewall re-moved and a rake used to measure captured mass flow.The top plate generates a boundary layer similar to thaton a vehicle forebody. The boundary-layer trips were steelballs with a 0.16 cm diam. Pitot and static pressure sur-veys were made at several vertical locations in the throatsof the side and center passages (Figs. 2 and 3) and at themass flow station; 102 static pressure orifices were distrib-uted throughout the model.

Mach 6 Experimental ResultsThe swept geometry inlet started easily at Mach 6 andsubsequent, recent data indicates enough spillage is pro-vided for starting below Mach 2.5. The Mach 6, theoreti-cal shock diagram in a horizontal plane was verified ex-

T O P S U R F A C E

S I D E P A S S A G E C E N T E R P A S S A G EFig. 3 Throat Mach number contours, mass weighted M3.0 (not to scale).

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- F L I G H T ^TUNNEL" A/ U - 6lDA TAM A S S W E I G H T E D D A T A ( M -6.0)

l.0r M A S S C A P T U R E R A T I O

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C O N T R A C T I O N R A T I OT H R O A T M A C H NO .R E C O V E R YnKM A S S C A P T U R E

P A S S A G ES I D E5. 93.0.4 6.968.3 2

C T R .7 .53.1.6 6.983.63

T OT ALINLET7. 03.1.5 9. 9 78.9 5

Fig. 4 Integrated performance parameters.

perimentally; the sidewall shocks corresponded to a totalturning of 6.8, including a 1.2 boundary-layer correction.The strut portions upstream of the throats shortened theinlet by furnishing 75% of the over-all inlet compression.

The survey data were processed and analyzed by a digi-tal computer program which by a curve-fitting interpola-tion procedure expanded the data into a network of ap-proximately 1000 grid points, covering the entire throatflow area. Besides performing numerical integrations,Mach number, total pressure, and unit mass flow werecalculated for each grid point; and contour maps of eachparameter were plotted by the computer's graphics sys-tem. The resulting Mach number contours are given inFig. 3 where the width scale is 7 times the height scale forboth passages.The side passage throat contours are relatively symmet-rical, and the predictions of boundary-layer thickness 5agree well for the top and side surfaces. A nearly horizon-

tal shock of about 8 turning was generated by the cowlleading edge and is still near the cowl surface at thethroat. This discrete shock is smeared by the interpolationprocess; and consequently a vertical Mach number gradi-ent, extending well beyond the predicted 5 for the cowl, isindicated. There is some rounding of the contours at thecorners, but no flow separation was detected. Note thatmost of the vertical walls were relieved near the cowl sur-face to counteract the excess compression produced by thecowl shock. The Mach number contours of the center pas-sage are not as symmetrical due to the greater shock waveconcentration, Fig. 2; however, the Mach 3 (mass weight-ed average) contour encloses the major portion of the totalarea.Integrated parameter performance is given in Fig. 4, forMach 6 flow in front of the inlet, which simulates a flightMach number of 7.6. The measured mass flow capture of95 % matches the predicted curve value. The curve isslightly higher than found in Ref . 4 because the fuel injec-tion struts and cowl have been moved upstream 4.1 cm toachieve the shock diagram of Fig. 2. The measured kineticenergy efficiency is 97.8% compared with the predictedtunnel value of 98.5%. The predicted curves are slightlyhigh primarily because the cowl shock and viscous cornerinteractions were not included. The aerodynamic contrac-tion ratio, which is based on the average throat Machnumber and total pressure, is 7.0 instead of the value of8.7 found in Ref. 4; because moving the struts upstreamincreased the throat width somewhat and because themeasured total pressure recovery was somewhat lower.The side passage total pressure recovery was lower thanthe center passage recovery because of the relativelythicker boundary layers in the side passage. The strut po-sitioning within the inlet provided a split in mass flow be-tween the center and side passages in the ratio of 63/32.


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SEPTEMBER 1974 E N G I N E E R I N G NOTES 591The small differences between the experimental and pre-dicted inlet performance will not have a significant effecton the engine thrust performance quoted in Ref. 4.

ReferencesiRenry, J. R. and McLellan, C. H., "The Air-BreathingLaunch Vehicle for Earth-Orbit ShuttleNew Technology and

Development Approach," Journal of Aircraft, Vol. 8, No. 5, May1971, pp. 381-387.2Becker, J. V., "New Approaches to Hypersonic Aircraft." Pre-sented at the 7th Congress of the International Council of theAeronautical Sciences, Rome, Italy, Sept. 1970; also Becker, J.V., "Prospects for Actively Cooled Hypersonic Transports," As-tronautics &Aeronautics, Vol. 9, No. 8, Aug. 1971, pp. 32-39.3Brown, D. A. et al., "Development of Liquid-Hydrogen Scram-jet Key to Hypersonic Flight," Aviation Week and Space Tech-nology, Sept. 17, 1973, pp. 75-78.4Henry, J. R. and Anderson, G. Y., "Design Considerations forthe Airf rame-Integra ted Scramjet," TM X-2895, 1973, N A S A .

Influence of Flaps and Engineson Aircraft Wake Vortices

David C. Burnham* and Thomas E. SullivantU.S. Department of Transportation, TransportationSystems Center, Cambridge, Mass.

ALTHOUGH previous investigations have shown that thenature of aircraft wake vortices depends on the aircrafttype and f lap configura t ion, the causes for these differ-ences have not been clearly identified. In this Note weshow that observed differences in vortex core structure arerelated to engine placement, engine thrust and wing f lapdeflection angle.Much of the quantitative in forma t ion on the velocitydistribution within aircraft vortices has been collected bythe Federal Aviation Administration's National AviationFacilities Experimental Center ( NAFE C ) in AtlanticCity, N.J. In these experiments (conducted at NAFECand Idaho Falls, Idaho) flight paths of test aircraft wereselected so the vortices would drift through an instru-mented smoke tower. Hot wire anemometers were used tomeasure vortex velocities. Smoke grenades placed at regu-lar intervals on the tower provided flow visualization ofthe vortex. These tests have shown that most vortices canbe divided into two general classes:

11) Tubular (T): The vortex has high tangential veloci-ties concentrated in a very tight core. Long tubular smokestreamers can be observed along the vortex axis whensuch a vortex passes near a smoke grenade on the tower.2) Nontubular (NT): The vortex has substantially lowertangential velocities and a large diffused vortex core. Lit-tle axial transport of injected smoke is observed.The types of vortices observed in NAFEC tower tests arelisted in Table 1. For all configura t ions , the engine thrustwas adjusted to maintain level flight past the tower . Thevortices from aircraft with four-wing mounted enginesReceived May 20, 1974. We would like to acknowledge the co-operation and assistance of NAFEC personnel and the ai rc raf t pi-

lots dur ing the tower fly-by tests reported here.Index categories: Aircraf t Aerodynamics ( Inc luding ComponentAerodynamics) ; Jets, Wakes, and Viscid-Inviscid Flow Interac-tions.*Staff Scientist.tStaff Engineer.

G R O U N DFig. 1 Geometry for acoustic scattering from aircraftvortices.

Table 1 Vortex typea vs aircraft type andconfigurationAi r c r a f t conf igura t ion6

Airc raf t type Holding Takeof f L an d i n gTTNT

Propeller driven (DC- 7 ) T TNo wing-mounted engines (B-727) T TFour wing-mounted engines (B-707) T Tc

a According to instrumented tower measurements (T = tubu-lar, NT = n o n tu b u la r ) .6 Flap extensions: Holding, none ; Takeof f , p a r t i a l ; Landing ,full.c Semitubular.configured fo r takeoff were designated "semitubular"(somewhat larger cores than in holding configura t ion).Note that the only significant differences occur in landingconfiguration.An independent but consistent vortex classification canbe obtained by interpreting data obtained with a pulsedbistatic acoustic vortex sensing system developed at theTransportation Systems Center.2 In this system acousticpulses ar e transmitted from one side of an aircraft flightpath and received on the other. The presence of a receivedsignal from the vortex depends on its acoustic ray-bendingproperties. The m a x i m u m scattering angle Qm (see Fig. 1)is particularly sensitive to the type of vortex core (for agiven circulation, the smaller the vortex core, the largerthe max i mu m scattering angle, 6m). Thus a tubular vor-tex would be expected to have a significantly larger valueof 6m than a nontubular vortex. Tests conducted at sever-al airports have shown that vortices from landing aircraftcould be classified on the basis of observed 6m with thesame results as in Table 1. Propeller driven aircraf t , air-craft with no wing-mounted engines and aircraf t with tw owing-mounted engines (DC-10, B-737), were found to givetypical values of 9m = 1.2 rad or higher. Aircraft withfour wing-mounted engines (DC-8 and B-707) typicallygave values of Qm = 0.5 rad. Intermediate values of 0m,which appeared to depend on the ambient wind condi-tions, were observed for the B-747 (four wing-mounted en-gines). Acoustic measurements made at N A F E C showthat aircraf t with four wing-mounted engines generatevortices with large scattering angles (6m ^ 1.0 rad) inboth holding and takeoff configurations.In order to explain the observed differences in corestructure one must take into account the effect of flapangle on the origin of a vortex from an aircraft wing. Ingeneral, the vortex core is generated at the edge of the liftdistribution, which in "holding" or "cruise" configuration(zero flap angle) is located at the tip. However, in landingconfiguration (full flaps) relatively little lift is generatedby that portion of the wing beyond the outboard edge ofthe flaps. The vortex generated by the strong lift disconti-nuity at the flap edge is therefore likely to dominate theformation of the vortex core with relatively little pertur-

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Carl A. Trexler- Performance of an Inlet for an Integrated Scramjet Concept