1、NASA TECHNICAL MEMORANDUM Ln LOAN COPY RETURN AFWL (4LIL-2) KIRTLAND AFB, N MEX JET EFFECTS ON BOATTAIL PRESSURE DRAG OF ISOLATED EJECTOR NOZZLES AT MACH NUMBERS FROM 0.60 TO 1.47 by Douglas E. Harrington Lewis Research Center Cleveland, Ohio WASHINGTON, Is fill bt MAY 1969 4SWNATIONAL AERONAUTICS A
2、ND SPACE ADMINISTRATIONProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NASA TM X-1785 TECH LIBRARY KAFBI NM I 0151416 JET EFFECTS ON BOATTAIL PRESSURE DRAG OF ISOLATED EJECTOR NOZZLES AT MACH NUMBERS FROM 0.60 TO 1.47By Douglas E. Harrington Lewis Re
3、search CenterCleveland, Ohio NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sole by the Clearinghouse for Federal Scientific and Technical InformationSpringfield, Virginia 22151 - CFSTI price $3.00Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-AB
4、STRACT The cylindrical ejector nozzles were operated over a range of pressure ratios from approximately 1.0 (jet off) to 11. Results were obtained with both 15 0 and 100 boattail angles. The 15 configurations utilized different radii of curvature at the boattail junc-ture with a cylindrical forebody
5、. Subsonically, the jet caused large reductions in boat-tail pressure drag whether the jet was under or overexpanded. Supersonically, however, reductions in boattail drag were obtained only if the jet was near full expansion or was underexpanded. A jet boundary simulator was effective in duplicating
6、 a fully expanded jet with an exit static-pressure ratio of one.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-JET EFFECTS ON BOATTAIL PRESSURE DRAG OF ISOLATED EJECTOR NOZZLES AT MACH NUMBERS FROM 0.6010 1.47 by Douglas E. HarringtonLewis Research
7、CenterSUMMARY An experimental investigation has been conducted to determine the effects of a cold jet on the boattail pressure drag of four isolated cylindrical ejector nozzles. The Mach number range was from 0.60 to 1.47. Nozzle pressure ratio was varied from approxi-mately 1.0 (jet off) to 11. The
8、 effects of secondary airflow were also studied. The noz-zle configurations included three with a 150 trailing-edge boattail angle and one with a 100 boattail. The boattail juncture with the cylindrical portion of the nacelle for the 15 configurations was faired with different radii of curvature. In
9、 addition, jet effects were simulated by a cylinder positioned downstream of the nozzle exit for the 15 configura-tions. At subsonic speeds, the jet caused large reductions in drag of the 15 boattails. This drag reduction was relatively insensitive to nozzle pressure ratio for values much less than
10、the design value. However, boattail drag was further reduced as the jet pres-sure ratio was increased to the design condition and beyond, thereby increasing the jet-exit static-pressure ratio and hence the tendency for jet pluming to occur downstream of the nozzle exit. Supersonically, the boattail
11、pressure drag was unaffected by the jet until it also approached full expansion. As it became underexpanded, the boattail drag was significantly reduced. The trends were basically the same for the 100 boattails except that boattail drag was affected to a lesser degree by the jet. In general, the eff
12、ect of increasing secondary flow was to decrease boattail pres-sure drag by increasing the jet-exit static-pressure ratio. Secondary flow was most effective in reducing boattail pressure drag coefficient at subsonic speeds when the nozzle was operating at or near full expansion or was underexpanded.
13、 A cylindrical jet bound-ary simulator was effective in duplicating a fully expanded jet with an exit to local am-bient static pressure ratio of one.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INTRODUCTION Current airbreathing propulsion systems
14、designed for supersonic flight operate over a wide range of nozzle pressure ratios. To maintain efficient operation at all flight speeds, variations in the nozzle expansion ratio are required. At subsonic speeds, for example, the exit area of a variable flap divergent ejector will be smaller than th
15、at re-quired at supersonic speeds. This reduction in exit area necessitates increased boat-tailing of the a.fterbody. The resultant drag can be a significant portion of the net thrust of the propulsion system, particularly at subsonic cruise where the engine is at a re-duced power setting. In additi
16、on, the jet issuing from the exit of the nozzle will have a pronounced effect on boattail drag (refs. 1 to 3). As part of a program in airbreathing propulsion at the Lewis Research Center, vari-ous nozzle concepts, designed primarily for supersonic cruise application, are being studied at off-design
17、 conditions. Subsonic and transonic performance is being obtained with cold-flow models in isolated nacelles in the Lewis 8- by 6-Foot Supersonic Wind Tunnel. These results will be compared with the installed performance of the same noz-zles obtained during flight tests using an F-106B aircraft. Nac
18、elles that house an after-burning J-85/13 turbojet engine as a gas generator will be installed under the large delta wing of the F-106B with the nozzles extending downstream of the trailing edge. Scale models of the F-106B are also being studied in the wind tunnel (ref. 4) to determine test-ing proc
19、edures that provide correlation with flight data. An experimental investigation, therefore, was conducted in the Lewis 8- by 6-Foot Supersonic V nd Tunnel to determine the effects of a cold jet on the drag of four isolated ejector nozzle. The Mach number range was from 0. 60 to 1.47 and nozzle-press
20、ure ratio was ed from approximately 1.0 (jet off) to 11. In addition, a cylinder was posi-tioned downit.ream of the nozzle exit for the 150 configurations to determine the effective-ness of a jet : oundary simulator in duplicating the effects of a jet on boattail pressure drag.SYMBOLS A cross-sectio
21、nal area CD pressure drag coefficient, D/qA C pressure coefficient, (p - p0)/q0 D drag d diameter F nozzle gross thrustProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-F - D - nozzle efficiency F+F5 F _D nozzle gross thrust coefficient M Mach number P
22、 total pressure secondary total pressure measured beneath primary nozzle actuating ring P secondary total pressure measured at station 7 p static pressure q dynamic pressure r boattail juncture radius of curvature T total temperature w weight flow rate w/ - 1 i_.scorrected secondary weight flow rati
23、o wpyTp v velocity x axial distance downstream of adapter-afterbody interface y distance measu-ed along primary rake from primary airflow passage wall zradial distance from model surface primary flap angle o boundary-layer Vickness momentum thickness Subscripts:i ideal m model p primary air s second
24、ary air 10 local ambient 0 boattail surface3-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-0 free-stream 7 nozzle inlet station 8 nozzle throat station 9 nozzle exit stationAPPARATUS AND PROCEDUREInstallation The nozzles were strut mounted in the t
25、est section of the Lewis 8- by 6-Foot Super-sonic Wind Tunnel as shown in figure 1. The geometry of the model and its thrust-measuring system are shown in figure 2. The main part of the model was a strut-supported cylinder with an ogive nose. The model external shell was grounded and was supported f
26、rom the tunnel ceiling by a hollow, vertical strut. The adapter portion of the model was attached to the air bottle, which was cantilevered by flow tubes from supply manifolds located outside the test section. Front and rear bearings supported the air bottle. Thus, the axial force acting on the floa
27、ting part of the model, including both the adapter and nozzle sections, was transmitted to the load cell, located in the nose of the model shell. The nozzle performance presented herein does not include the friction drag measured on the floating portion of the model designated as the adapter section
28、. The downstream end of the adapter section was arbitrarily selected as being 0. 75 model diameter upstream of the nozzle throat (station 8). Friction drag on the adapter section was estimated using the semiempirical, flat-plate, local skin-friction coefficient (given in fig. 6 of ref. 5) as a funct
29、ion of free-stream Mach number and Reynolds number. The coefficient accounts for variations in boundary-layer thickness and flow profile with Reynolds number. Previous measurements of the boundary-layer characteristics at the aft end of this jet exit model in the 8- by 6-Foot Supersonic Wind Tunnel
30、indicated that the profile and thickness were essentially the same as that computed for a flat plate of equal length. The strut wake appeared to affect only a localized region near the top of the model and resulted in a lower local free-stream velocity than measured on the side and bottom of the mod
31、el. Therefore, the results of reference 5 were used without cor-rection for three-dimensional flow effects or strut interference effects. The calculated friction drag of the adapter section was, therefore, added to the load cell reading to obtain the thrust-minus-drag of the nozzle section. A static
32、 calibration of the thrust-measuring system was obtained by applying a known force to the nozzle and measuring the output of the load cell. To minimize changes in the calibration due to variations in temperature (e.g., aerodynamic heating due to exter-nal flow), the load cell was surrounded by a wat
33、er-cooled jacket and was maintained at a constant temperature. 4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Primary and secondary air were provided by means of airflow supply lines which entered the model through the support strut. Secondary air
34、in the central air bottle passed through crossover struts inside the model to simulate cooling flow for the primary nozzle and the internal surface of the outer shroud. A uniform primary flow was main-tamed by using a choke plate and two straightening screens upstream of station 7. Primary weight fl
35、ow rate was determined from static and total pressure and temper-ature measurements at station 7 and a calibration constant for the flow system based on measurements for a standard ASME nozzle. Secondary weight flow rate was determined from a standard ASME flow-metering orifice located in the second
36、ary air-supply line. The ambient pressure was constant for a given free-stream Mach number; thus, a vari-ation in nozzle pressure ratio was obtained by varying the nozzle total pressure P7. Nozzle Geometry The four nozzle configurations that were tested are shown in figure 3. Three of these configur
37、ations had 150 boattail trailing-edge angles. Since the emphasis of the test was on the external drag rather than on internal performance, simple cylindrical ejectors were Used for the internal flow. The nozzle exit diameter was 5. 725 inches (14. 542 cm). The boattail juncture with the cylindrical
38、portion of the nacelle for these 150 configura-tions was smoothed with different radii of curvature. These radii were 0 (sharp edge), 0. 5, and 2. 5 model diameters. In order to simulate another trailing-edge flap position, a 100 boattail angle with a radius of curvature of 0. 5 model diameter was a
39、lso tested. The 100 nozzle was the cylindrical-ejector type, but it had an exit diameter of 6.636 inches (16. 855 cm). In conjunction with the F- 106B flight tests at Lewis Research Center, a General Electric J_85/13 primary nozzle was simulated for this test. Since the primary nozzle of the J-85/13
40、 engine has a variable throat, two different throat areas were used with each of the four afterbody configurations in this test. The smaller throat area corresponded to the minimum reheat setting (Primary I) while the larger area sim-ulated maximum reheat (Primary II). The actuating mechanism for va
41、rying primary throat area was simulated by a ring containing 12 slots. Secondary air was forced to flow beneath this ring by means of a deflector. In addition to these configurations, the nozzles with 15 boattails were tested in the presence of a cylindrical section extending from the nozzle exit, (
42、station 9) as shown in figure 3(e). The purpose of the cylinder was to approximate the local flow field that would exist if a jet were fully expanded with an exit static-pressure ratio p9/p1 0 of 1. 0. The jet boundary simulator diameter was equal to the nozzle base diameter.5Provided by IHSNot for
43、ResaleNo reproduction or networking permitted without license from IHS-,-,-Instrumentation Afterbody pressure instrumentation for atypical configuration is shown in figure 4(a). It was assumed that the flow field over the afterbody was symmetric about the strut cen-terline. Thus, with the exception
44、of one row of instrumentation, all boattail static-pressure orifices were located on one side of the afterbody. Limitations in space neces- sitated one of the rows being at 3000 instead of 600. Boattail static-pressure orifices were located at the centroids of equal projected areas and the pressure-
45、drag coefficient was computed using the method described in reference 6. Extra pressure taps were located just downstream of the boattail juncture of the 150 (r/dm = 0) configuration to help define boattail pressure distributions. These pressures were located at 00 , 900, and 1800 and were not used
46、for pressure-drag deter-mination. In addition, three rows of static-pressure orifices were located on the cylin- drical portion of each nozzle at 00, 90, and 1800 . The axial location of each afterbody orifice was determined by the distance x downstream of the adapter-nozzle interface. Table I gives
47、 the position of each orifice as well as a nondimensional position coordinate x/dm . Nozzle exit pressure p9 was assumed to be the average of five pressures located in the exit plane (station 9). Local ambient pressure P1 0 was assumed to be the aver-age of the six pressures located at the trailing
48、edge of the boattail and in the same axial plane as orifice 13. Details of temperature and pressure instrumentation at station 7 are shown in fig-ure 4(b). Pressures in the primary airflow passage were measured by two static-pressure orifices and a total-pressure rake containing 11 tubes. Primary no
49、zzle total pressure was obtained from an integrated average of these pressures. The accompany-ing table lists pressure orifice spacing as distance y from the inner surface of the pas-sage. Secondary-air total pressure P was measured using four total-pressure tubes. Primary- and secondary-air total temperatures were measured by copper -con stantan thermocou
