NASA-TN-D-3021-1965 Flight-determined low-speed lift and drag characteristics of the lightweight m2-f1 lifting body《轻量级的m2-f1上升机身的飞行测定的低速升力和阻力特性》.pdf

《NASA-TN-D-3021-1965 Flight-determined low-speed lift and drag characteristics of the lightweight m2-f1 lifting body《轻量级的m2-f1上升机身的飞行测定的低速升力和阻力特性》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-3021-1965 Flight-determined low-speed lift and drag characteristics of the lightweight m2-f1 lifting body《轻量级的m2-f1上升机身的飞行测定的低速升力和阻力特性》.pdf（45页珍藏版）》请在麦多课文档分享上搜索。

1、PASA c N 0 T n z c 4 r/) 4 z TECHNICAL NOTE NASA TN D-3021 _z- 0 0 FLIGHT-DETERMINED LOW-SPEED LIFT AND DRAG CHARACTERISTICS OF THE LIGHTWEIGHT M2-FI LIFTING BODY - by Victor W. Horton, Richard C. Eldredge, and Richard E. Klein Flight Research Center Edwards, Gal$ NATIONAL AERONAUTICS AND SPACE ADMI

2、NISTRATION WASHINGTON, D. C. SEPTEMBER 1965 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FLIGHT-DETERMINED LOW- SPEED LIFT AND DRAG CHARACTERISTICS OF THE LIGHTWEIGHT M2-F1 LIFTING BODY By Victor W. Horton, Richard C. Eldredge, and Richard E. Klei

3、n Flight Research Center Edwards, Calif. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $2.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IH

4、S-,-,-FLIGHT-DETERMINED LOW-SPEED LIFT AND DRAG CHARACTERISTICS OF THE LIGHTWEIGHT I-Fl LIFTING BODY By Victor W. Horton, Richard C. Eldredge, and Richard E. Klein Flight Research Center SUMMARY The low-speed lift and drag characteristics of a manned, lightweight M-2 lifting-body vehicle were determ

5、ined in unpowered free-flight tests at angles of attack from 0“ to 22O (0.38 radian) and at calibrated airspeeds from 61 knots to 113 knots (31.38 to 58.13 meters/second). pared with results from full-scale wind-tunnel tests of the same vehicle. Flight data are com- The investigation showed that 95

6、percent of the vehicle maximum lift-drag ratio of 2.8 was available through an angle-of-attack range from 4.4“ to 14.6“ (0.08 to 0.25 radian). be low in comparison with most other aircraft, no serious difficulties were experienced in landing the test vehicle. Although this lift-drag ratio is conside

7、red to The lift and trim characteristics were linear in the angle-of-attack range from 0“ to 15“ (0.26 radian). Although the same vehicle was tested in flight and in the wind tunnel, significant differences existed in the values of zero-lift drag and drag due to lift. INTRODUCTION In recent years, m

8、any wind-tunnel studies have been made during the development of lifting reentry configurations capable of gliding to a speci- fied recovery site and making a conventional horizontal landing. To comple- ment these studies, the NASA Flight Research Center conducted exploratory flight tests of the M-2

9、 lifting-body vehicle. The M-2 configuration was selected for the flight investigation because of the relatively large amount of aerodynamic data available for the vehicle from previous wind-tunnel studies (refs. 1 to 6). A lightweight version of the M-2 was chosen because of the advantages offered

10、in design simplicity, low cost of construction, simple manual operation of the controls, and ease of maintenance, modification, and repair. This approach also enabled flight Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-data to be obtained within a

11、 relatively short time. A glider-type operation was adopted in preference to on-board propulsion in order to simplify the design and construction of the vehicle and to avoid possible uncertainties in the effects of power on vehicle performance, stability, and control. This paper presents the low-spe

12、ed lift and drag characteristics deter- mined in flight for the lightweight M-2 configuration, designated the W-Fl, and compares flight data with full-scale wind-tunnel-test results for the same vehicle. In addition, the rather unusual construction and flight-test techniques used in the program are

13、discussed. The flight tests were conducted at the Flight Research Center, Edwards, Calif., at altitudes below 13,000 feet (3962 meters) and at calibrated airspeeds from 61 knots to ll3 knots (31.38 to 58.13 meters/second). SYMBOLS Physical quantities used in this paper are given, where applicable, i

14、n both the U.S. Customary Units and in the International System of Units (SI). Factors relating the two systems are presented in reference 7. “2 longitudinal acceleration, g an normal-acceleration factor (ratio of the net aerodynamic force along the airplane Z-axis to the weight of the airplane), g

15、D drag coefficient, - (2s CD CL cP % CIS base drag coefficient, - L lift coefficient, - qs AP pressure coefficient, - qc D drag force along flight path, pounds (kilograms) d 65 distance flown for test, average true speed x time, feet (meters) gravitational ace elerati on, fee t/s econd2 (meter s/s e

16、cond2 ) Ah corrected altitude loss, feet (meters) h altitude energy condition, feet (meters) measured altitude loss, feet (meters) pressure altitude, feet (meters) E AND FLIGHT OPERClTIONS M2-Fl Hull and internal structure.- The M2-Fl is comprised of two major The hull assembly is constructed of 3/3

17、2-inch assemblies: the hull, which includes the cockpit and control surfaces, and the internal structure. (2.38 mm) mahogany plywood skin and 1/8-inch (3.18 mm) mahogany rib sections reinforced with spruce. The exterior surface is wrapped with Dacron and doped to provide a more durable finish. The v

18、ertical fins, rudders, and elevons are thick slab sections constructed with 0.016-inch (0.41 mm) aluminum skin. The trailing-edge flaps are composed of welded 0.028-inch (0.71 mm) aluminum tubing covered with Dacron and are equipped with fixed trim tabs to reduce the stick forces to a comfortable le

19、vel. Turning vanes were attached to the side of the hull (for flight and wind-tunnel tests) to alleviate the flap “buzz“ in the 95-knot to lO5-knot (48.87 to 34.02 meters/second) speed range, which was a result, apparently, of the vortex pattern shed from the vehicle base. A modified glider canopy o

20、f molded Plexiglas and plywood encloses the cockpit and access hole that was provided for removal of the internal structure. A Plexiglas nose and side window are also included to provide additional visibility during landing. Styrofoam tail skids were placed on the hull to prevent damage in the event

21、 of overrotation in landing or takeoff. The internal structure (fig. 16), which is constructed of welded steel tubing, includes the fixed landing-gear assembly, control stick, rudder pedals, and control system from the cockpit to a mixer plate. The nosewheel and main-wheel assemblies are slightly mo

22、dified light-aircraft types, and the main-gear shock and strut assembly incorporates both a viscous damper and a bungee. The seat is a modified rocket ejection seat. Differential main- wheel braking and a steerable nosewheel are provided for ground control. Control system.- The control system is con

23、ventional. Gearing ratios were initially determined with the a.id of a ground-based simulation of the vehicle response characteristics, then adjusted to be conventional with flight experiences of the pilot (ref. 13) . consists of both a trailing-edge flap and elevons. For pitch control, the elevons

24、are deflected approximately 2.2 times the flap deflection, as shown in figure 10. This ratio results in the elevons maintaining a nearly constant 10 (0.17 radian) local angle of,attack at all trimmed conditions (alocal = 1.7a + Se). flection of the elevons, and directional control is provided by the

25、 rudders. The longitudinal forces are reduced from an estimated 23 pounds to 30 pounds (13.34 to 13.61 kilograms) pull force to about 5 pounds to 10 pounds (2.27 to 4.54 kilograms) by means of fixed tabs on the flaps. and elevon forces are very light and require the use of bungees to provide the des

26、ired “feel. “ The longitudinal-control system Roll control is obtained through differential de- The rudder 11 . - . . .- - . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Landing-assist rocket.- Because of the anticipated low lift-drag ratio and po

27、or visibility from the cockpit of the M2-Fl during the flare portion of the landing, some means was deemed necessary to provide the pilot with extra time for maneuvering in the event of a difficult landing. A simple and efficient means for increasing the flare time was to equip the vehicle with a sm

28、all solid-propellant rocket motor with the thrust vector alined longitu- dinally through the center of gravity. The rocket, when fired, effectively increases the maximum lift-drag ratio from 2.8 to 4.5. The rocket, a small solid-propellant type developed by the Naval Ordnance Test Station at China L

29、ake, Calif., was installed in the base of the E-Fl. It provided a nominal thrust of 180 pounds (801 newtons) for approximately 11 seconds. Flight Operations The flight program for the M2-Fl began with a series of taxi tests on car tow to check out the control rigging and to familiarize the pilot wit

30、h the ground stability of the vehicle. experience, tow speeds were gradually increased (lift-off was achieved at about 75 KCAS (38.78 m/sec), until a maximum of about lo5 KCAS (54.02 ,/see) was reached, which corresponded to 87 percent of the design limit speed. About 60 airborne car tows were compl

31、eted before the first airplane tow was made. As the pilot acquired confidence and The light wing loading and unknown control and stability of the test vehicle presented a possible problem, in that the pilot could lose control of the vehicle if the turbulence in the wake of the tow plane were encount

32、ered. To determine an acceptable range of tow positions and to assess the effects of takeoff acceleration, several trial tows were made using a conventional sail- plane. The results of these tests indicated that a high tow position, about l5O feet (45.7 meters) above the C-47 airplane, and a towline

33、 length of about 1000 feet (304.8 meters) would minimize the wake effects of the tow airplane. Also, before the first airplane tow, four rocket firings were made with the M2-F1 to demonstrate the structural integrity of the rocket installation and to discover any possible adverse thrust effects on t

34、he stability and con- trol of the vehicle. Two of the firings were made with the vehicle in motion. The first test, during a car tow at about 60 KCAS (30.87 ,/see) with only the nosewheel off the ground, revealed no noticeable pitch or yaw perturbations. The second firing, also during a car tow, was

35、 made while the vehicle was air- borne at an altitude of approximately 10 feet (3.05 meters) and a speed of 95 KCAS (48.87 m/sec) after towline release. believed the vehicle stability to be slightly improved while the rocket was burning. In this instance, the pilot Although available to the pilots o

36、n all flights, the landing-assist rocket was used only twice during the program of 37 flights. In one instance, the rocket was used as a precautionary measure when turbulence was encountered during the flare portion of the flight. had leveled off too high and used the rocket to insure a low vertical

37、 velocity at touchdown. In both cases, the pilot made a normal landing and reported In the other, the pilot felt that he 12 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-that the rocket was beneficial in increasing the apparent lift-drag ratio and

38、the lateral-directional stability. All air-tow tests were made in the early morning to take advantage of the normally calm air. face winds exceeded 5 knots (2.57 meters/second). experience, this requirement was relaxed until flights were made in steady, 10-knot to 15-knot (5.14 to 7.22 meters/second

39、) winds with light turbulence. Initially, the flights were postponed if steady sur- As the pilot acquired more Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-APPENDIX B SPEC UL CALIBRATIONS It was impractical to position the airspeed hea.d far enoug

40、h in front of the vehicle body to eliminate flow-interference effects on the static-pressure source and angle-of-attack vane. Detailed calibrations of the angle-of-attack and airspeed systems were, therefore, required. Angle-of-Attack Calibration Angle of attack was measured by a small vane attached

41、 to the nose boom and was calibrated during level, unaccelerated airplane-towed flight. The vane readings were compared with those from the longitudinal accelerometer and the recording inclinometer. 0 = y +a, which, for the assumed flight conditions, reduces to y = 0 and 0 = a. Both instruments were

42、 alined within 1 minute of arc (0.00029 radian) with the top surface of the vehicle, which was used as a reference for angle of attack. The inclinometer, therefore, recorded angle of-attack directly; whereas, the accelerometer readout was related to angle of attack by means of the expression 0 = a =

43、 sin-la2. Results from both the flight and wind-tunnel calibrations are shown in figure 5. The entire angle-of-attack range avail- able for flight could not be calibrated in flight because of minimum tow-plane speed at the higher angles and the M2-Fl towed structural speed at the lower angles. The w

44、ind-tunnel tests, on the other hand, covered a large angle-of- attack range and agreed well with the flight calibration. This technique makes use of the relationship Airspeed Calibration During the wind-tunnel tests, airspeed data were obtained corresponding to four trimmed angles of attack in fligh

45、t. These points, shown as squares in the calibration presented in figure 6, lie essentially along a straight line. Flight calibration points were obtained by ground towing at the lower speeds and by air towing beside a pacer airplane at the higher speeds. A 2.5-statute-mile (4023.4 meters) course on

46、 Rogers Dry Lake was used for the ground tows over which the stabilized indicated airspeed and elapsed time were recorded for traverses in both directions in order to compensate for any wind effects. temperature, and average ground speed, the calibrated airspeed was determined. The corresponding ind

47、icated airspeed was corrected for instrument error, which was obtained from a laboratory calibration of the instrument. Calibration data obtained by this method (triangular symbols in fig. 6) were limited by the minimum safe lift-off speed of the M2-Fl and the maximum speed obtainable by the tow car

48、. By taking into account the test altitude, ambient 14 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The pacer calibration was obtained during airplane tow by taking data points simultaneously from the M2-F1 and the pacer airplane, after both were

49、stabilized in formation, at a series of indicated airspeeds. The calibration data, using this method, were limited by the same factors as the angle-of- attack calibration. The calibrations obtained by these three techniques showed excellent agreement. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-APPENDIX C T