1、RESULTS OF A BRIEF FLIGHT INVESTIGATION OF A COIN-TYPE STOL AIRCRAFT by Terrell W. Feistel und Robert C. Innis Ames Reseurch Center Moffett Field Cui$ , -_ NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. AUGUST 1967 i Provided by IHSNot for ResaleNo reproduction or networking permitt
2、ed without license from IHS-,-,-TECH LIBRARY KAFB. NM NASA TN D-4141 RESULTS OF A BRIEF FLIGHT INVESTIGATION OF A COIN-TYPE STOL AIRCRAFT By Terrell W. Feistel and Robert C. Innis Ames Research Center Moffett Field, Calif. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sole by the Clearinghouse f
3、or Federal Scientific ond Technical Information Springfield, Virginia 22151 - CFSTI price $3.00 I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RESULTS OF A BRIEF FLIGHT INVESTIGATION OF A COIN-TYPE STOL AIRCRAFT By Terrell W. Feistel and Robert C.
4、 Innis Ames Research Center SUMMARY The airplane tested to gain experience with a COIN (for Counter INsurgency) type STOL aircraft had two propellers driven by turbine engines, and double-hinged, single-slotted flaps to deflect the slipstream on the largely immersed wing. It was capable of good low-
5、speed performance and had acceptable handling qualities in the STOL regime (with landing and take-off distances consistently less than 800 feet over a 50-foot obstacle), provided the possibility of engine failure was ignored. This performance was achieved, despite flaps with only medium effectivenes
6、s, because the aircraft had a low aspect ratio, a high power loading, and a “no-flare“ landing gear design. It is shown (for the sake of an interesting comparison) that the performance of the test aircraft, as flown (ignoring engine failure), compared favorably with that of a large four-engined STOL
7、 aircraft, tested previously, which was much more sophisticated (it included a fail-safe propulsion system). above the minimum single-engine control speed, however, in compliance with the normal safety restrictions for twin-engine airplanes, major aspects of the performance of the test aircraft are
8、no better than that obtainable with many small “twins“ in current production and most of the original objectives. of the COIN concept are compromised. If flown INTRODUCTION The NASA has for several years studied COIN-type STOL aircraft with primary emphasis on STOL operational problems, low-speed ha
9、ndling qualities, the desirability of propeller interconnect (cross shafting), and high-lift devices. Recently, in extensive wind-tunnel tests, NASA studied various small-scale COIN models (ref. 1) to determine low-speed stability and control problems and the effect of configuration changes. Results
10、 of simulator studies (ref. 2) and flight tests of the Ryan VZ-3 were used to analyze handling qualities of COIN aircraft. The first flying prototype COIN, the Convair Model 48 “Charger,“ was flight tested for 10 hours to examine the problems further and to obtain additional operational experience w
11、ith STOL aircraft. The results of these flight tests which are considered to be pertinent to a general understanding of COIN-type (i.e., relatively small, simple, and inexpensive) STOL aircraft are presented here. Provided by IHSNot for ResaleNo reproduction or networking permitted without license f
12、rom IHS-,-,-SYMBOLS A “X CL CD CF FEa 6, - - F f h hi P R C R D R S Re - - - - S SHP SL SW TED TEU t 2 actual geometric aspect ratio axial acceleration, g lift coefficient net drag coefficient (power on), drag coefficient minus thrust mean equivalent skin-friction coefficient in cruise lateral stick
13、 force, lb longitudinal stick force, lb - cz Dcruise a equivalent flat plate area, sq ft = CD S cruise ground clearance, ft indicated pressure altitude, ft roll rate, deg/sec rate of climb, fpm rate of descent, fpm rate of sink, fps mean Reynolds number surface length and reference wing area, shaft
14、horsepower - - in cruise, 2 x (where 2 is the mean wetted v is the kinematic viscosity of air) sq ft total landing distance over 50 feet, ft total aircraft wetted area, sq ft trailing edge down trailing edge up thrust coefficient, time thrust qs Provided by IHSNot for ResaleNo reproduction or networ
15、king permitted without license from IHS-,-,-vki MC W S ai P - Y A % 6F ssp stab 0 6 0 cp cp indicated airspeed, knots minimum single-engine control speed, knots wing loading, psf indicated angle of attack, deg sideslip angle, deg approach descent angle, deg elevator tab deflection, deg flap deflecti
16、on angle, deg spoiler deflection, deg stabilator deflection, deg pitch attitude, deg pitch rate, deg/sec p itch ac c e lera t ion , rad/s ec2 roll angle, deg rolling acceleration, rad/sec2 DESCRIPTION OF AIRCRAFT AND INSTRUMENTATION Description of the Test Aircraft The test aircraft (shown in fig. 1
17、 in an STOL landing approach) had two 650 SHP engines driving opposite rotation, 9-foot-diameter propellers; the tips rotated upward in the center. Retractable Krueger flaps were used on the inboard leading edges of the wing which was largely immersed in the propeller slipstream. slotted and double
18、hinged, and was deflected 60/30 for the landing approach and 2Oo/O0 for take-off. 40/110 position is an intermediate one which was investigated only briefly. ) The control system was entirely mechanical; lateral control was obtained with circular-arc spoilers only and longitudinal control with a fre
19、e-floating, single-hinged (geared, camber-changing) horizontal tail (called a stabilator). The twin rudders were conventional. A three-view drawing of the test airplane is shown in figure 2. Table I lists the pertinent physical characteristics. The 44-percent chord trailing-edge flap was single (The
20、 flap geometry is shown on the inset; the 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Design Considerations The Charger was designed according to a Marine Corps specific operating requirement (SOR) that specified a take-off and landing distance
21、 of 500 feet over a 50-foot obstacle and included a requirement for “single-engine sur- vivability.“ To meet this latter requirement, a “torque-equalizer device was incorporated to reduce the power on one engine automatically in the event the other failed, thereby allowing the pilot to hold the wing
22、s near-level long enough to eject safely; this device was operable during all the NASA flight tests in the STOL regime (with the exception, of course, of the single engine investigations). Instrumentation The flight test data were recorded with an on-board tape deck and photo These data were correla
23、ted through voice contact with the pilot by a panel. time-coding system. Measured quantities included: airspeed, altitude, angle of attack, angle of sideslip, rate of climb, and engine data on the photo panel; plus angular rates (about three axes), angular accelerations (about three axes), linear ac
24、celerations (in three directions), and primary control positions recorded on tape. For the landing and take-off data, a “TOL“ (Take-Off and Landing) camera (ref. 3) was installed in the bottom of the fuselage to provide accurate information on aircraft height (ground clearance), ground speed, and pi
25、tch attitude. mSULTS AND DISCUSSION The flight test data were obtained during nine flights encompassing approximately 10 hours of flying time. In two later flights, landing and take-off data were obtained with the “TOL“ camera. were removed to allow the camera a full field of view. All of the take-o
26、ffs and landings and most of the NASA flights were made in the STOL regime (i.e., at speeds below power -off stall and substantially below the minimum single- engine control speed). each covering a particular category of flight: first, landing; second, take- off; third, cruise; and fourth, miscellan
27、eous areas (engine-out, wave-off, etc.). Most of the parts have separate sections dealing with performance, handling qualities , and operational techniques and pilot s comments . The aft landing-gear doors The discussion of the data is divided into four parts, STOL Landing Regime Most of the NASA fl
28、ights were made in the landing configuration because the landing maneuver has proven to be the prime problem area for most STOL aircraft; also, the landing distance of this aircraft, as with most projected aircraft in the STOL category, is consistently greater than that for take-off. 4 Provided by I
29、HSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Landing performance.- Figure 3 shows the flight-test-measured STOL landing distances over 50 feet for the test airplane (corrected for the small winds) as a function of descent angle; air distance is also indicated,
30、 and data are shown from both NASA and contractor flight tests. Representative Breguet 941 landing distances are also shown for comparison; frequent refer- ence will be made to this aircraft, a four-propeller deflected-slipstream STOL transport which was flight tested by NASA in 1963 (ref. 4). The c
31、urves super- imposed on the data show calculated distances for approaching over a 50-foot obstacle in a straight-line descent, touching down with no flare, and stopping with an average deceleration (a) of 1/2 g. These calculated distances will be used in subsequent plots to show the probable trend o
32、f landing distance variation as approach angle and speed are varied. angle of distance over 50 feet is noteworthy, since it points out the great importance of attaining good descent performance, with low speed, for short -f ield landings. The variation with descent The permissible sink speed of 16 f
33、ps for the prototype landing gear made possible no-flare landings with a great reduction in landing distance and an improvement in touchdown point consistency over the conventional full flare. While investigating the landing performance, the NASA pilot was concerned primarily with different combinat
34、ions of approach speed and sink rate and not with optimizing the ground roll (as evidenced by the increased ground distance at the higher descent angles caused by bouncing). tractor flight testing, actual landing distances as low as 600 feet over 50 feet were achieved. shown in the figure. ) During
35、the earlier con- (For reference, the contractor landing data are also Descent characteristics.- The descent characteristics in the landing approach condition, derived (after cross-plotting and extrapolation) from the NASA flight test data, are presented in figure 4 as a function of speed, for consta
36、nt throttle settings, with angle-of-attack values superimposed. The actual descent capability (ignoring the possibility of engine failure) was determined by a stall margin of approximately 10 knots and 10 angle of attack usually demanded by the pilot (the indicated angle of attack for the stall vari
37、ed from approximately 20 to 35O as power was increased from idle to NRP). The nominal approach condition for a STOL landing was made at 55 to 60 knots with 600 to 800 fpm rate of descent (corresponding to slightly more than half the NRP); this was approximately 10 knots below the power-off stall spe
38、ed of 67 knots. wing loading of 37-40 psf). The weight ranged from 7100 to 7700 pounds (corresponding to a The nominal descent condition for the Breguet 941 (ref. 4) is also indicated and is approximately the same (in terms of speed, CL, and c) as that shown for the subject airplane. conditions for
39、these two fairly dissimilar craft-is interesting. The fact that the Breguet descends at the same angle as the test aircraft, despite its higher aspect ratio (6-1/2 vs. 4-1/2), is attributable to its superior flap effectiveness. landing gear had not been drup-tested, a 16-fps placard was imposed. The
40、 similarity of the descent - The design sink speed for the gear was 20 fps, but since the prototype 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Figure 5 shows the actual power-on, lift-drag polars of the airplane as derived from flight test dat
41、a. Lines of constant thrust coefficient are plotted with angle-of-attack values superimposed. The nominal STOL approach condition is indicated, and the Breguet 941 condition is again shown for %ax comparison. The figure points up the large disparity between the capability of the airplane and the act
42、ual CL which is representative of most STOL designs. to the steep descent capability required for short landing distances over an obstacle and partly to the stall margin usually demanded by the pilot (see ref. 5 for further discussion of this subject). used in a landing approach, This disparity is d
43、ue partly Comparison between flight tests.- Constant STOL approach condition - descent capabilities determined in wind-tunnel and thrust coefficient (Tc = 1.0) polars in the nominal are plotted in figure 6. Data derived from the NASA flight-tests are compared with data from tests in the Ames 40- by
44、80-oot Wind Tunnel (ref. 6) with a full-scale COIN model (similar to the test air- plane but with blunt wing tips). Data are also shown for a 1/4-scale model similar in configuration to the test airplane (and corrected to the same configuration as that tested in the 40- by 80-foot tunnel). the plot
45、for reference are calculated no-flare landing distance lines for a wing loading of 40 psf. Superimposed on The full-scale wind-tunnel data agree surprisingly well with the flight T, = 1.0) test data; the small-scale data, however, appear optimistic and might encourage the prediction of a nominal lan
46、ding distance (even for of 500 to 600 feet, as opposed to the TOO to 800 feet which is actually feasible. This illustrates a common general weakness in small-scale STOL low- speed data, that is, an optimistic (rather than conservative, as might be expected, because of the low Reynolds number) indica
47、tion of descent capability from small-scale wind-tunnel test data. to attempt to explain this discrepancy, but the significance is obvious when the effect on the preliminary design is considered. It is beyond the scope of this paper Handling qualities. - Table I1 shows the most important stability,
48、con- trol, and handling qualities parameters, with some appropriate pilot ratings, for the test airplane in the nominal STOL approach condition.2 were obtained from stabilized descents made in the landing configuration at altitudes of 2,000 to 10,000 feet. The handling qualities characteristics in t
49、his condition were generally acceptable, for flight test purposes at least. Some of the more noteworthy aspects will be discussed below. These data Control System Characteristics The problems encountered with the nonboosted control system were not precisely as might have been anticipated. stabilizer-elevator system, controlled by a small trailing-edge tab connected dire
