NASA-TN-D-8242-1976 Aerodynamic characteristics of a vane flow angularity sensor system capable of measuring flightpath accelerations for the Mach number range from 0 40 to 2 54《当马.pdf

1、AERODYNAMIC CHARACTERISTICS OF A VANE FLOW ANGULARITY SENSOR SYSTEM CAPABLE OF MEASURING FLIGHTPATH ACCELERATIONS FOR THE MACH NUMBER RANGE FROM 0.40 TO 2.54 Glenn M. Sakamoto Dryden Flight Research Center Edwards, Cali$ 93523 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. MAY 1976

2、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM I 111111 Ill11 Ill11 lllll lllll lllll lllll llll I - _ - _._ . . -. -_- 1. Report Nn NASA TN D-8242 4. Title and Subtitle MRODYNAMIC CHARACTERISTICS OF A VANE FLOW ANGULARITY SENS

3、OR SYSTEM CAPABLE OF MEASURING FLIGHTPATH 4CCELERATIONS FOR THE MACH NUMBER RANGE FROM 0.40 TO 2.54 7. Authorts) Glenn M . Sakamoto I 2. Government Accession No, I - -_ - I - _ -. -. . _ . - -. - - . . - 9. Performing Organization Name and Address NASA Flight Research Center P.O. Box 273 Edwards, Ca

4、lifornia 93523 2. Sponsoring Agency Name and Address National Aeronautics and Space Administration Washington, D .C . 20546 5. Supplementary Notes 0334008 - 3. Recipients Catalog No. 5. Report Date May 1976 6. Performing Organization Code 8. Performing Organization Report No. H-900 - 10. Work Unit N

5、o. 517-51-01 - 11. Contract or Grant No. Technical Note - 13. Type of Report and Period Covered 14. Sponsoring Agency Code _ 6. Abstract This report presents the results of a wind tunnel inves- tigation of the aerodynamic characteristics of a vane flow angularity sensor system capable of measuring f

6、lightpath accelerations . The aerodynamic characteristics of the angle of attack vane and the angle of sideslip vane are summarized. The test conditions ranged in free stream Mach number from 0.40 to 2.54, in angle of attack from -2O to 22O, in angle of sideslip from -2O to 12O, and in Reynolds numb

7、er from 5.9 X 10 per meter (1.8 X 10 per foot) to 18.0 X 10 per meter (5.5 X 10 per foot) . 6 6 6 6 The results of the wind tunnel investigation are compared with results obtained with similar vane configurations. Com- parisons with a NACA vane configuration are also made. In addition, wind tunnel-d

8、erived upwash for the test installation is compared with analytical predictions. 7. Key Words (Suggested by Author(s) ) Wind tunnel tests Vane flow direction sensors Upwash corrections - - 18. Distribution Statement Unclassified - Unlimited Category: 06 21. NO. of Pages 22. Price 43 I $3.75 19. Secu

9、rity Classif. (of this report) 20. Security Classif. (of this page) Unclassified Unclassified *For sale by the National Technical Information Service, Springfield, Virginia 22161 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-AERODYNAMIC CHARACTERIS

10、TICS OF A VANE FLOW ANGULARITY SENSOR SYSTEM CAPABLE OF MEASURING FLIGHTPATH ACCELERATIONS FOR THE MACH NUMBER RANGE FROM 0.40 TO 2.54 Glenn M. Sakamoto Flight Research Center SUMMARY A wind tunnel investigation was conducted on a vane flow angularity sensor system capable of measuring flightpath ac

11、celerations in an attempt to define the aerodynamic characteristics of the vanes. The measurements of the angle of attack and angle of sideslip vanes both showed error throughout the Mach number range tested. The vane errors were greatly affected by Mach number. The errors due to Mach number were la

12、rgest at transonic Mach numbers for the angle of attack vanes. The errors were largest at high supersonic Mach numbers for the angle of sideslip vane and resulted from shock interference. The characteristics of the effects of Mach number on both vanes indicated that the vane assemblies were in such

13、close proximity to each other that there was con- siderable mutual interference. A comparison of the calibration obtained for this test installation with a calibration obtained for a similar system substantiates this finding. A comparison of the calibrations for the tested vanes and standard NACA va

14、nes showed their aerodynamic characteristics to be similar. This similarity was observed despite significant differences in the vane configurations. The frequency and damping of the tested vanes were one-half those of the NACA vanes. A comparison of wind tunnel-derived nose boom upwash with analytic

15、al pre- dictions showed that the prediction method provided reasonable estimates of upwash over the Mach number range tested. The prediction method did not account for the shock-to-vane interactions manifest in the wind tunnel data. The results of this calibration verify and extend previously obtain

16、ed wind tunnel results for a similar configuration. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INTRODUCTION Aircraft flight test programs require the precise measurement of angle of attack and angle of sideslip, since these measurements yield in

17、formation concerning the aircrafts flight status and are also essential to many of the subsequent analyses. One of the most common measurement systems is a nose boom-mounted vane system with a configuration similar to the NACA design described in references 1 to 3. Recently, flow direction vanes hav

18、e also been used as the mount for linear acceler- ometers, which are used to sense vehicle flightpath acceleration. Flow angularity vanes equipped with flightpath accelerometers were developed during a wind tunnel program conducted at the Arnold Engineering Development Center (AEDC) in Tullahoma, Te

19、nn . (refs. 4 to 6) and are significantly different in design from con- ventional NACA vanes. The new sensor system was used in a joint NASA/Air Force transonic aircraft technology (TACT) program at Edwards, Calif. This was NASAs first experience with a vane flow angularity sensor system capable of

20、measuring flightpath accelera- tions. Since unusual variations in angle of attack and angle of sideslip were observed in limited wind tunnel tests of a similar system (refs. 4 to 6), in 1973 the entire nose boom assembly for the TACT airplane was calibrated in the Unitary Plan Wind Tunnel facility a

21、t the NASA Ames Research Center. This report presents the results of the angle of attack and angle of sideslip vane calibration obtained at Ames. During the investigation, Mach number ranged from 0.40 to 2.54, angle of attack varied from -2O to 22O, and angle of sideslip varied from -2O to 12O. The

22、calibration data from these tests are compared with previous cali- brations obtained at AEDC . The calibration data are also compared with a calibration for a NACA vane configuration to identify areas of aerodynamic similarity. Finally, a comparison is made between the nose boom upwash predicted by

23、wind tunnel results and those predicted by an analytical prediction technique. SYMBOLS Physical quantities in this report are given in the International System of Units (SI) and parenthetically in U .S . Customary Units. The measurements were taken in Customary Units. Factors relating the two system

24、s are presented in reference 7. angle of attack vane error, a - a deg Ea i t angle of sideslip vane error, P 2 . cos a t - Pt, deg P E natural frequency, H z fn h altitude, m (ft) M Mach number 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-R R Re

25、/Z radial distance from the nose boom centerline to the vane center of pressure (fig. 16) , approximately 20 cm (7.87 in.) approximate radial distance from the nose boom centerline to the shock wave along the vane hinge line (fig. 16) , cm (in .) approximate radial distance from the nose boom center

26、line to the shock wave just inboard of the vane (fig. 16) cm (in.) unit Reynolds number, per m (per ft) components of velocity, V (fig. 24) , m/sec (ft/sec) velocity , m/sec (ft/sec) rectangular coordinates of the body axes (fig. 24) angle of attack, deg angle of sideslip , deg damping ratio angle o

27、f attack vane float, deg a -a it nose boom upwash, - i a angle of sideslip vane float deg P:* - P, L -I- nose boom sidewash , Pi“ Sub scripts : e equivalent i indicated t true m free stream Superscript: * transformed from nose boom axis to wind axis 3 Provided by IHSNot for ResaleNo reproduction or

28、networking permitted without license from IHS-,-,-TEST APPARATUS The subject of this investigation is a flow angularity sensor system consisting of a Pitot-static probe system for measuring static and total pressures, vane assemblies for measuring angle of attack and angle of sideslip, and a dual-ax

29、is accelerometer capable of measuring longitudinal and normal acceleration along a flightpath. The flow angularity sensor is shown in figure 1 and its dimensions are given in figure 2. Figure 1. Photograph of flow angularity sensor system. (0.75) FI ig htpat h accelerometer housing Plan view Pitot-s

30、tatic I probe system (6.27) 8.26 (3.25) 35.56 End view 23.42 T- A 23.42 (9.22) (9.22) (14.00) (9.28) Angle of sideslip vane Side view Figure 2. Dimensional details of flow angularity sensor in three views. Dimensions are in centimeters (inches). 4 Provided by IHSNot for ResaleNo reproduction or netw

31、orking permitted without license from IHS-,-,-The flow angularity vanes are flat plates with swept, beveled leading edges. The vanes are fixed to rotating , aerodynamically contoured posts. The dimensions of the vanes, which were identical for angle of attack and angle of sideslip , are shown in fig

32、ure 3. Further details concerning the design and selection of this vane configuration are given in reference 4. i 19105 (7.50) 2.24 (0.88) 1.27 (0.50) View A-A 2.36 4 I, (0.93) Figure 3. Dimensions are in centimeters (inches). Dimensional details of vanes in three views. The angle of attack vanes ar

33、e symmetrically arranged. Angle of attack measure- ments were taken with respect to the left vane as seen in normal plan view orientation. The angle of sideslip vane assembly is 23.57 centimeters (9.28 inches) to the rear of the angle of attack vane assembly. Because of the size of the flightpath ac

34、celerometer assembly, the angle of attack vanes for this particular system are physically constrained to deflections of -5O and 28O relative to the centerline of the housing. Figure 4 illustrates the arrangement of the flightpath accelerometer assembly. Figure 4. Schematic of flightpath acceleromete

35、r assembly. 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WIND TUNNEL TESTS Tunnel Description The flow angularity calibration was conducted in the 11-Foot Transonic and the 9- by 7-Foot Supersonic Wind Tunnels in the NASA Ames Unitary Plan Wind

36、Tunnel facility. Both are closed circuit, continuous flow, variable density tunnels. The transonic leg has a Mach number capability from 0.40 to 1.40, and the supersonic leg has a Mach number range from 1.55 to 2.55. Both tunnels are described in detail in reference 8. Static Tunnel Calibration The

37、strut-mounted nose boom assembly (fig. 5) together with the externally controlled drive unit were first calibrated under static conditions. The nose boom assembly was put in a level position and statically loaded in the angle of attack plane to determine nose boom and wind tunnel balance flexibility

38、 in the presence of aerodynamic lift. An upward aerodynamic load of approximately 445 newtons (100 pounds) , which was determined from an equation in reference 9, was applied. The effect of this load was determined to be negligible. No side loads were applied in the angle of sideslip plane. Flow ang

39、ularity sensor Location of angle of sideslip 8.26 10.49 (3.25) (4.13) vane centerline Nose boom Location of angle of attack vane centerline 23.57 (9.28) (29.50) 178.44 (70.25) 240.28 (94.60) Figure 5. Schematic of wind tunnel nose boom installation. Dimensions are in centimeters (inches). The extern

40、ally controlled drive unit was calibrated by hard mounting a laser unit to the nose boom and using it to trace a locus of the strut position on a fixed target located upstream of the test section. In addition, a split-bubble clinometer accurate to 6 seconds was mounted on the boom and simultaneous r

41、eadings were taken from it and the vane synchro transmitter. A bubble level accurate to 3 seconds was affixed to the angle of attack vane to indicate the level vane condition. Hysteresis was detected in the control drive unit and was minimized by always approaching the angle of attack set in the tun

42、nel from a higher angle of attack. 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Test Procedure The center of rotation in the angle of attack plane was at the angle of attack vane centerline such that at zero angle of sideslip the orientation of

43、the nose boom assem- bly was identical to the tunnel centerline. In the angle of sideslip plane, the center of rotation was in the strut body such that all nonzero angles of sideslip were off the tunnel centerline. Upright and inverted tests were performed to establish tunnel flow angularity correct

44、ions. The angle of attack set in the tunnel was always approached from a higher angle of attack to minimize mechanical and aerodynamic hysteresis. Schlieren photographs were taken at certain Mach number, angle of attack, and angle of sideslip conditions in both tunnels for shock wave visualization.

45、Emphasis was placed on the transonic Mach number region to define compressibility effects on the vanes and on Mach numbers near 2.0, where sideslip vane error reversal was known to occur (refs. 4 to 6). Test Conditions The wind tunnel test conditions are summarized in table 1. Mach number ranged fro

46、m a nominal 0.40 to 2.54. Angle of attack ranged from -2O to 22O in the 11-Foot Tunnel and from -2O to 16O in the 9- by 7-FOOt Tunnel. Angle of sideslip ranged from -2O to 12O in both tunnels. The angles of attack and angles of sideslip set in the tunnel were varied in 2O increments. Reynolds number

47、 ranged from 6 6 6 6 5.9 X 10 per meter (1.8 X 10 per foot) to 18.0 X 10 per meter (5.5 X 10 per foot). Several Reynolds numbers were tested at free stream Mach numbers of 0.90, 1.30, and 1.51 to assess the effects of unit Reynolds number. Instrumentation, Data Reduction, and Accuracy The instrument

48、ation already installed in the nose boom assembly was adequate for the flow angularity calibration. Flow angularity measurements were obtained from synchro transmitters in the angle of attack and angle of sideslip vanes. Simultaneous free stream total and static pressure measurements were obtained f

49、rom the Pitot-static probe system. The measurements were recorded on standard wind tunnel recorders. The flow angularity calibration data for the angle of attack and angle of sideslip vanes were corrected for the effects of wind tunnel flow angularity. Corrections for boom bending based on the analytical method described in reference 10 w