1、pANDNASA TECHNICAL NASA TM X-2917MEMORANDUM447I(IASA-c;ix-2917)- -I!ME5TIGA2I-61-oF .7 -2-06!,Z,EVEAAL 1ACA 1 SEJIES AXISY0LEI:C ILETSI.4ACH IZUIBEhS 2301 0.4 1. 29 (NASk)-3 p HC $4 75 CSCL 01A UnclasH1/01 36248AN INVESTIGATION OF SEVERALNACA 1-SERIES AXISYMMETRIC INLETSAT MACH NUMBERS FROM 0.4 TO 1
2、.29by Richard J. ReLangley Research CenterHampton, Va. 23665NATIONAL AERONAUTICS AND SPACE ADMINISTRATION * WASHINGTON, D. C. * MARCH 1974Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1. Report No. 2. Government Accession No. 3. Recipients Catalog
3、No.NASA TM X-29174. Title and Subtitle 5. Report DateAN INVESTIGATION OF SEVERAL NACA 1-SERIES AXISYM- March 1974METRIC INLETS AT MACH NUMBERS FROM 0.4 TO 1.29 6. Performing Organization Code7. Author(s) 8. Performing Organization Report No.Richard J. Re L-858810. Work Unit No.9. Performing Organiza
4、tion Name and Address 760-64-60-02NASA Langley Research Center 11. Contract or Grant No.Hampton, Va. 2366513. Type of Report and Period Covered12. Sponsoring Agency Name and Address Technical MemorandumNational Aeronautics and Space Administration 14. Sponsoring Agency CodeWashington, D.C. 2054615.
5、Supplementary Notes16. AbstractAn investigation was conducted in the Langley 16-foot transonic tunnel to determinethe performance of seven inlets having NACA 1-series contours and one inlet having anelliptical contour over a range of mass-flow ratios and at angle of attack. The inlet diam-eter ratio
6、 varied from 0.81 to 0.89; inlet length ratio varied from 0.75 to 1.25; and inter-nal contraction ratio varied from 1.009 to 1.093. Reynolds number based on inlet maximumdiameter varied from 3.4 x 106 at a Mach number of 0.4 to 5.6 X 106 at a Mach number of1.29.17. Key Words (Suggested by Author(s)
7、18. Distribution StatementInlet Unclassified - UnlimitedNACA 1-series inletsInlet performanceSTAR Category 0119. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price*Unclassified Unclassified 144 $4 , 7,For sale by the National Technical Information Serv
8、ice, Springfield, Virginia 22151.1,Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-AN INVESTIGATION OF SEVERAL NACA 1-SERIES AXISYMMETRICINLETS AT MACH NUMBERS FROM 0.4 TO 1.29By Richard J. ReLangley Research CenterSUMMARYAn investigation was conduct
9、ed in the Langley 16-foot transonic tunnel to determinethe performance of seven inlets having NACA 1-series contours and one inlet having anelliptical contour over a range of mass-flow ratios and at angle of attack. The inlet diam-eter ratio varied from 0.81 to 0.89; inlet length ratio varied from 0
10、.75 to 1.25; and inter-nal contraction ratio varied from 1.009 to 1.093. Reynolds number based on inlet maxi-mum diameter varied from 3.4 x 106 at a Mach number of 0.4 to 5.6 x 106 at a Machnumber of 1.29.INTRODUCTIONThe development of airfoil sections which delay the formation of strong shocksuntil
11、 high supercritical local Mach numbers are reached has potentially opened the wayfor aircraft to operate efficiently at higher subsonic Mach numbers. The airfoil sectioncharacteristics demonstrated in wind-tunnel tests (refs. 1 to 5) have been confirmed inflight tests on a straight-wing airplane (re
12、f. 6) and on a sweptback-wing airplane (refs. 7to 9).Evolution of complete transport aircraft configurations capable of cruising near aMach number of 1.0 requires development of engine nacelles with performance compatiblewith airframe capability if the full potential of the supercritical airfoil is
13、to be realized.Since the size and operating characteristics of an aircraft affect the number and locationof engine macelles, various aircraft configurations will likely be considered. However,it is probable that most, if not all, configurations will have turbofan engines with axisym-metric or pitot-
14、type inlets. Since the cruise speeds proposed for such transports aresubstantially above those of current subsonic transports, little of the inlet data in existencewould aid in the design of a suitable inlet. The most comprehensive investigations ofaxisymmetric inlets (NACA 1-series) were conducted
15、at low speeds and are reported inreferences 10 and 11. Investigations in the transonic speed range (refs. 12 to 17) wereconducted on inlets having diameter ratios that are small compared with those requiredProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-
16、,-,-for high-bypass-ratio turbofan engines. The results of reference 18 indicate that inletexternal pressures at high subsonic speeds are not significantly affected by the long cowlnacelle afterbody flow field. Therefore, valid evaluations of the external performanceof axisymmetric inlets can be mad
17、e experimentally without integrated development ofnacelle afterbody shape.The present investigation was conducted to obtain force and pressure data on sevenNACA 1-series inlets and one elliptical profile inlet at mass-flow ratios typical of thoseof high-bypass-ratio turbofan engines. Five of the NAC
18、A 1-series inlets had differentexternal shapes and two of them had greater amounts of internal area contraction. Inletdiameter and length ratios were in the ranges from 0.81 to 0.89 and 0.75 to 1.25, respec-tively. Internal area contraction ratio varied from 1.009 to 1.093.The investigation was cond
19、ucted in the Langley 16-foot transonic tunnel at Machnumbers from 0.4 to 1.29 and at angles of attack from 00 to about 60 at selected Machnumbers. Reynolds numbers based on model maximum diameter ranged from about3.4 x 106 at a Mach number of 0.4 to 5.6 x 106 at a Mach number of 1.29.SYMBOLSA area n
20、ormal to inlet center lineAxial forceCA axial-force coefficient,qxAmaxCA,E external axial-force coefficient, External axial forceqooAmax1 AmaxCA, F forebody pressure-force coefficient, 1 A m Cp dAAmax AspCA,w axial-force coefficient due to skin friction on rake support struts, and ductwall and cente
21、rbody surface between rake face and structural breakCp pressure coefficient, -q0D diameterd intake diameter of NACA 1-series inlet (difference between Dh and twicethe inlet lip radius)M Mach number2Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-fm m
22、ass flowmf/mfoo inlet mass-flow ratio, 1 rVrpoAhV Prrp static pressurePt stagnation pressureq dynamic pressureR radius measured from model center lineR free-stream Reynolds number based on maximum diameter of modelr lip radiusTt free-stream stagnation temperatureV velocityX length of inlet from lip
23、to start of cylindrical part of modelx distance from lip of inlet measured longitudinallyY maximum ordinate measured perpendicular to reference line at maximum-diameter station for NACA 1-series inletsy local ordinate measured perpendicular to reference line for NACA 1-seriesinletaangle of attack wi
24、th respect to model center line, degp densitymeridian angle, measured from top of model in clockwise direction whenlooking upstream, degL-8588 3Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Subscripts:b force balancec force balance cavitycr critica
25、l condition corresponding to local sonic flowD point at which CA,E reaches 1.1 times the level of CA,E at lowerMach numbersh most forward point on inlet lip1 localmax maximummin minimumr mass-flow rake station in ductse between seal and external surface of modelsi between duct wall and sealsp stagna
26、tion point on inlet lipw duct wall at mass-flow rake station0o free-stream conditionMODELThe model had a maximum diameter of 45.72 cm and was mounted in the test sectionby a rear sting. The inlet part of the model was supported by a force balance and wasstructurally isolated from the afterbody which
27、 was attached directly to the sting. Photo-graphs showing the model installed in the wind-tunnel test section are presented as fig-ure 1 and a simplified cross-sectional sketch of the model assembly with the shortestinlet (NACA 1-85-75) is shown in figure 2.4Provided by IHSNot for ResaleNo reproduct
28、ion or networking permitted without license from IHS-,-,-Eight inlets were designed, seven of which had NACA 1-series outer profiles andthe eighth inlet had an elliptical outer profile. Three of the NACA 1-series inlets hadthe same outer profile but different amounts of internal area contraction. Th
29、e nondimen-sional NACA 1-series outer profile ordinates as presented for a given lip radius in ref-erence 10 are reproduced in table I. A summary of the important geometric parametersfor each inlet is contained in figure 3. The elliptical inlet is referred to herein by thedesignation “Elliptical-85-
30、100“ as a matter of convenience so that its geometric desig-nation is analogous to the NACA 1-series inlets. Its external coordinates can be computedfor an ellipse having a major axis of 91.440 cm and a minor axis of 3.348 cm. Nondimen-sional outer profile radii measured on a precision measuring mac
31、hine are presented intable II and internal design ordinates for each inlet are presented in table m. The inter-nal ordinates for the NACA 1-85-100 inlets with contraction ratios of 1.046 and 1.093 areelliptical between the lip leading edge and minimum duct area station (throat). From thethroat to th
32、e 25 percent station, the internal contour of all the inlets consisted of a 10semicone expansion. The remainder of the internal contour consisted of a faired curvewith a maximum slope of 6.30 for the shortest inlet. The proportional rate of area growthas a function of distance in the faired section
33、was identical for all inlets.Static-pressure orifices were drilled into tubing placed in grooves in the modelsurface and covered with a filler material. The longitudinal locations of the orifices oneach inlet outer profile forward of the structural break (station 69.85) are presented intable IV. The
34、 four struts which connected the inlets to the centerbody were used to routethe inlet static-pressure tubes to differential pressure-scanning units mounted in the noseof the centerbody. Three of the struts were instrumented with the pressure probes neces-sary to measure the duct mass flow. (See fig.
35、 4.) The model was constructed mainly ofaluminum with some of the primary structure such as the sting made of steel. The struc-tural break between the force-balance-mounted inlet and the sting-mounted afterbodyhoused a flexible seal strip to prevent airflow through the gap. (See detail sketch infig.
36、 2.)For convenience, the model aft of the structural break is considered to be the after-body ( 1.36Dmax in length) although a part of the model (0.53Dmax in length) mounted onthe force balance is cylindrical. The ordinates and orifice locations for the afterbodypresented in table V are based on the
37、 medium length inlet (X = 45.72 cm), where inletlength is only that part of the model defined by the inlet designation. The afterbody wasattached directly to the sting by means of four struts as shown in figure 2. All pressuretubes associated with the afterbody were routed through the four supportin
38、g struts intothe sting and out through the tunnel support system to individual differential pressuretransducers or differential pressure scanning units.5Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The mass-flow throttle plug was driven by an inte
39、rnally housed remote-controlledmotor and had a travel of about 25.4 cm aft of the position shown in figure 2. The openarea at the exit of the model (normal to the free-stream flow direction) was varied from678.55 cm2 to 1229.59 cm2 with the plug in its two extreme positions.WIND TUNNELThe investigat
40、ion was conducted in the Langley 16-foot transonic tunnel which is asingle-return atmospheric wind tunnel with continuous air exchange. The test section isoctagonal in shape with 4.724 meters between opposite walls (equivalent to the area of acircle 4.85 meters in diameter) and has axial slots at th
41、e wall vertices. The total widthof the eight slots in the vicinity of the model is approximately 3.7 percent of the test-section perimeter. At Mach numbers from 1.2 to 1.3, the divergence angle of the test-section walls is adjusted (based on calibration data) as a function of airstream dewpointtempe
42、rature to eliminate longitudinal static-pressure gradients that would occur on thecenter line because of condensation of atmospheric moisture. The solid blockage of themodel in the test section is between 0.88 percent (no flow through model) and 0.33 percent(throttle plug area only).The tunnel sting
43、 support system pivots in such a manner that the model remains onor near the test-section center line through the angle-of-attack range.TESTS AND METHODSEach inlet was tested at Mach numbers from 0.4 to 1.2 or 1.29 at an angle of attackof 00 and over a nominal angle-of-attack range from 00 to 60 at
44、Mach numbers of 0.4, 0.8,0.9, and 0.98 subject to force balance load limitations. At an angle of attack of 00, datawere taken at mass-flow ratio increments of 0.1 at all Mach numbers except 0.90, 0.96,and 0.98 where the increment was 0.05. At angle of attack, data were taken only at themaximum mass-
45、flow ratio obtainable (throttle plug in its most aft position). Sketchesshowing the variations in inlet geometry included in this investigation are shown at thetop of figure 3.The variations of free-stream stagnation temperature and Reynolds number (basedon maximum model diameter) with Mach number a
46、re shown in figure 5. For all the datapresented herein, boundary-layer transition on the inlets was not artificially fixed sincethe Reynolds numbers at high subsonic speeds were approximately 50 percent of typicalflight values at altitudes approaching the tropopause. A limited amount of data obtaine
47、dwith boundary-layer transition fixed on the external surface (0.125-cm-wide strip of6Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-number 100 silicon carbide particles, 2.54 cm aft of lip) showed no effect when comparedwith the free transition pre
48、ssure data.Angle of attack has been corrected for deflection of the sting and balance due tomodel aerodynamic forces and moments and for tunnel stream angularity. Duct massflow was calculated by using the rake-area-weighted stagnation pressure measurements(fig. 4) and static pressures on the rake, c
49、enterbody, and duct walls. The axial-forcedata were adjusted to the condition of free-stream static pressure at the structural breakstation between the balance-supported inlet and the sting-supported afterbody parts of themodel. That is, separate corrections were made in the area external to the seal in thebreak in the outer wall and in the area internal to the
