1、I I , NASA TECHNICAL NASA IM X-3324 MEMORANDUM I . tf . CII M rn AN INVESTIGATION OF SEVERAL NACA 1-SERIES INLETS AT MACH NUMBERS FROM 0.4 TO 1.29 FOR MASS-FLOW RATIOS NEAR 1.0 Richard J. Re Laagley Research center NATIONAL AERONAUTICS AND SPACE ARMIAIISTRATION * WASHINGTOM, D. C. DtcEMsER 19P5 Prov
2、ided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Richard J. Re NASA Langley Research Center Hampton, Va. 23665 National Aeronautics and Space Administration An investigation to determine the performance of eight NACA 1-series inlets at mass- flow ratios n
3、ear 1.0 was conducted in the Langley 16-foot transonic tunnel. The inlet diam- eter ratios (ratio of inlet diameter to maximum diameter) were 0.85 and 0.89 for an inlet length ratio (ratio of inlet length to maximum diameter) of 1.0. Inlet lip radius varied from 0.061 cm to 0.251 cm, and internal co
4、ntraction area ratio (ratio of inlet area to throat area) varied from 1.006 to 1.201. Reynolds number based on model maximum diameter ranged from 3.6 X lo6 at a Mach number of 0.4 to 5.9 x 106 at a Mach number of 1.29. The results indicate that nearly uniform pressure distributions on a given inlet
5、were obtained over a limited range of mass-flow ratios and Mach numbers. When inlet lip thick- ness was increased by means of lip radius or contraction ratio, the inlet critical Mach num- ber decreased. Drag-divergence Mach number inferred from forebody pressure integrations was above 0.94 for most
6、of the inlets tested. Unclassified - Unlimited NACA 1-series inlets Inlet performance For sale by the National Technical Information Service, Springfield, Virginia 221 61 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-AN INVESTIGATION OF SEVERAL NAC
7、A 1 -SERIES INLETS AT MACH NUMBERS FROM 0.4 TO 1.29 FOR MASS-FLOW RATIOS NEAR 1.0 Richard J. Re Langley Research Center An investigation to determine the performance of eight NACA 1-series inlets at mass-flow ratios near 1.0 was conducted in the Eangley 16-foot transonic tunnel. The inlet diameter r
8、atios (ratio of inlet diameter to maximum diameter) were 0.85 and 0.89 for an inlet length ratio (ratio of inlet length to maxi-mum diameter) of 1.0. Inlet lip radius varied from 0.061 cm to 0.251 cm, and internal contraction area ratio (ratio of inlet area to throat area) varied from 1.006 to 1.201
9、. Reynolds number based on model maximum diameter ranged from 3.6 X lo6 at a Mach number of 0.4 to 5.9 X lo6 at a Mach number of 1.29. The results indicate that nearly uniform pressure distributions on a given inlet were obtained over a limited range of mass-flow ratios and Mach numbers. Wen inlet l
10、ip thickness was increased by means of lip radius or contraction ratio, the inlet critical Mach number decreased. Drag-divergence Mach number inferred from forebody pressure integrations was above 0.94 for most of the inlets tested. INTRODUCTION The development of airfoil sections which delay the fo
11、rmation of strong shocks until high supercritical local Mach numbers are reached has opened the way for the design of lifting surfaces for efficient transport aircraft in the high subsonic speed range. Section characteristics of a super critical airfoil determined in a wind tunnel are pre- sented in
12、 references 1 to 5. Flight verification of wind-tunnel airfoil section character- istics for unswept and sweptback wings of finite span are contained in references 6 to 10. The development of turbofan engines of various sizes and bypass ratios and the advent of the supercritical airfoil section prov
13、ide the airplane designer with sufficient tools to design a wide variety of cruise-efficient subsonic transport airplanes. No one airplane configuration would satisfy the variety of performance requirements possible in this speed range. However, many configurations would probably incorporate turbofa
14、n Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-engines with air induction systems using axisymmetric or pitot-type inlets. These inlets, because of the high mass flows of turbofan engines, would have large diameter ratios (ratio of inlet diameter
15、to nlaximum diameter). In addition, some configurations with high subsonic cruise speeds require these inlets to operate at supercritical speeds; little high-speed inlet data exists, however, to aid in nacelle design for advanced sub- sonic transports. The most comprehensive data on axisymmetric inl
16、ets (NACA 1 -series) (reported in refs. 11 and 12) were obtained at low speeds. Investigations in the transonic speed range (refs. 13 to 18) were conducted on inlets which, in comparison with inlets required for most turbofan engines, had small diameter ratios. To complement the data just mentioned,
17、 several NACA 1 -series inlets of large diameter ratio were investigated over a range of mass flows in the Langley 16-foot transonic tunnel at Mach numbers from 0.4 to 1.29. These data are reported in reference 19. This investigation extends the external pressure-distribution results for four of the
18、 inlets discussed in reference 19 to higher inlet mass-flow ratios and includes external pressure distributions for four additional inlets. Mass-flow ratios near 1.0 were obtained using a model afterbody with a larger throttleable exit area than was used in the previous investigation. All of the inl
19、ets investigated here had a length ratio (ratio of inlet length to maximum diameter) of 1.0 and diameter ratios of 0.85 to 0.89. Inlet lip radius varied from 0.061 cm to 0.251 cm and internal contraction ratio (ratio of inlet area to throat area) varied from 1.006 to 1.201. The investigation was con
20、ducted in the Langley 16-foot transonic tunnel at a O0 angle of attack over a range of mass-flow ratios and at small angles of attack for mass-flow ratios near 1.0. Reynolds number based on mdel maximum diameter ranged from 3.6 x 106 at a Mach number of 0.4 to 5.9 X lo6 at a Mach number of 1.29. SYM
21、BOLS A area normal to inlet center line %,F integrated forebody axial-force coefficient (positive downstream), cpdA PQ - p, pressure coefficient, q, intake diameter of NACA 1-series inlet (difference between Dh and twice inlet lip radius) Provided by IHSNot for ResaleNo reproduction or networking pe
22、rmitted without license from IHS-,-,-diameter Mach number inlet mass-flow ratio, ,) prvrdA Q,AhVw static pressure dynamic pressure radius measured from model center line free-stream Reynolds number based on maximum diameter of model lip radius stagnation temperature velocity length of inlet from lip
23、 to start of cylindrical part of model, X = 45.72 em distance from lip of inlet measured longitudinally maximum ordinate measured perpendicular to reference line at maxinum diameter station for NACA 1 -series inlets local ordinate measured perpendicular to reference line for NACA 1-series inlet angl
24、e of attack with respect to model center line, deg density meridian angle, measured from top of model in clockwise direction when looking upstream, deg Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Subscripts: er critical condition corresponding to
25、 local sonic flow d duct D point at which external axial-force coefficient reaches 1.1 times its low Mach number level (from force-balance data of ref* 19) h most forward point on inlet lip Q local rnax maximum min minimum P point at which CA reaches peak (negative) value r mass-flow rake station in
26、 duct I S P stagnation point on inlet lip w duct wall at mass-flow rake station 00 free-stream condition MODEL The model consisted of an inlet and afterbody and had a maximum diameter of 45.72 em. The model was mounted in the wind-tunnel test section by a rear sting. A simplified cross-sectional ske
27、tch of the model assembly with an NACA 1-85-100 inlet is shown in figure 1. Eight NACA 1-series inlets (45.72 cm in length) were used for this investigation. Four of the inlets were used in the investigation described in reference 19 and four additional inlets were constructed for the present invest
28、igation. The variations in inlet geometry include inlet diameter ratio, inlet lip radius, and inlet internal area contrac- tion ratio. The nondimensional NACA I-series outer profile ordinates, as presented for a given Pip radius in reference 11, are reproctuced in table I. Figure 2 contains a Provid
29、ed by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-summary of the important geometric parameters for each of the inlets. Nondimension- alized internal ordinates for each inlet are shown in table II. The inlets with internal contraction ratios of 1.046 or grea
30、ter are elliptical between the lip and the minimum duct-area station (throat). From the throat to the 25-percent station (X/X = 0.25)1, the internal contour of all the inlets consisted of a lo semicone expansion. A faired curve made up the remainder of the internal contour. The proportional rate of
31、area growth (based on the difference between maximum duct area and throat area) as a function of distance in the faired section was Identical for all inlets. Diffuser area ratios (ratio of maximum duct internal area to inlet throat area) for each inlet are shown in table 11. Static-pressure orifices
32、 were drilled into tubing placed in grooves in the model surface and which had been covered with a filler material. The locations of the orifices on each inlet outer profile are presented in table III. The four struts which connected the inlets to the centerbody were used to route the inlet static-p
33、ressure tubes to differ- ential pressure-scanning units mounted in the nose of the centerbody. Three of the struts were instrumented with the pressure probes necessary to measure duct mass flow (see fig. 3). The inlets and afterbody were made from aluminum; parts of the pri- mary structure, the stin
34、g, for example, were made of steel. The afterbdy had a cylind ical external shape (2.89 in length) and was 45.72 em in diameter. The afterbody had a constant internal diameter back to sta- tion 111.76 where the duct changes to a conical shape, and thereafter increased in diam- eter to the exit. Exte
35、rnal orifice locations for the afterbody are presented in table W; the locations are based on inlet length (where inlet length X = 45.72 em). The after- body was attached directly to the sting by four struts (see fig. 1). The mass-flow throttle plug was driven by an internally housed remote-control
36、electric motor and had a travel of about 25.4 em. The open area at the exit of the model (normal to the free-stream flow direction) could be varied from 1022.42 ern2 to 1573.46 cm2 with the plug in either of its two extreme positions. WIND TUNNEL The investigation was conducted in the Langley 16-foo
37、t transonic tunnel, a single- return atmospheric wind tunnel with continuous air exchange. The test section is oetag- onal in shape measuring 4.724 m between opposite walls (an area equivalent to a circle 4.85 m in diameter). The test section has axial slots at the wall vertices; the total width of
38、the eight slots in the vicinity of the model is approximately 3.7 percent of the test sec- tion perimeter. At Mach numbers from 1.2 to 1.3, the divergence angle of the test sec- tion walls was adjusted (based on calibration data) as a function of airstream dewpoint temperature; the adjustment elimin
39、ated longitudinal static-pressure gradients that would Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-have occurred on the center line because of condensation of atmospheric moisture. A complete description of the wind tunnel and its airflow charact
40、eristics is contained in reference 20. The solid blockage of the model in the test section is between 0.88 per- cent (no flow through model) and 0.33 percent (throttle plug area only). The tunnel sting-support system pivots in such a manner that the model remains on or near the test- section center
41、line through the angle-of-attack range. TESTS AND METHODS Each inlet was tested at Mach numbers from 0.40 to 1.01 at an angle of attack of 00. The NACA 1-89-100 inlet with a contraction ratio of 1.006 was also tested at Mach numbers of 1.20 and 1.29. Most of the inlets were also tested at a nominal
42、angle of attack between 10 and 2O with a high mass-flow ratio at several subsonic Mach numbers. Sketches show- ing the variations in inlet geometry included in this investigation are presented in figure 2. The variations of free-stream stagnation temperature and Reynolds number (based on maximum mod
43、el diameter) with Mach number are shown in figure 4. All the data pre- sented here are for free boundary-layer transition on the model since the tunnel stream conditions and the large scale of the model result in a Reynolds number approaching one- half that of a nacelle 2.13 m in diameter at operati
44、ng altitude. Inlet angle of attack was obtained by correcting the angle of the model support sys- tem for deflection of the sting under aerodynamic loads and for test-section stream angu- larity. No corrections were made to the pressure data for test-section wall-interference effects or for local co
45、ndensation effects that may have occurred in the model flow field. The presence of the mass-flow throttle plug at the base of the afterbody affects the after- body pressure field; therefore, the small amount of afterbody pressure data presented should be considered qualitative. PRESENTATION OF RESUL
46、TS The results of this investigation are presented in graphic form as model external surface-pressure coefficients in figures 5 to 21. External and internal pressure coeffi- cients were machine plotted as a function of nondimensionalized inlet length for each mass-flow ratio and were faired by conne
47、cting adjacent readings with straight line seg- ments. Critical pressure coefficients are shown in figures 6 to 13 for reference pur- poses. The pressure-coefficient data for each inlet are presented in the figures as follows: Provided by IHSNot for ResaleNo reproduction or networking permitted with
48、out license from IHS-,-,-Inlet designation NACA 1-85-100 + NACA 1-89-100 NACA 1-89-100 Lip radius, cm Figures showing Internal contraction pressure coefficients for - ratio, A/A, C Figure 22 presents the variation of forebody axial-force coefficient (numerical integration in axial direction of exter
49、nal surface-pressure force coefficients from the stagnation point on the lip, as determined from the pressure distributions, to the rnaxi- mum diameter) with Mach number. This figure includes the appropriate data from reference 19 at lower mass-flow ratios. Figure 23 shows inlet lower critical Mach nurn- ber as obtained from cross plotting peak negativ
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