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  1. SOME CORRELATIONS OF FLIGHT-MEASURED AND WIND-TUNNEL MEASURED STABILITY AND CONTROL CHARACTERISTICS OF HIGH-SPEED AIRPLANES
    Authors: Walter C. Williams, Hubert M. Drake and Jack Fischel
    Report Number: NACA-H56AG62
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Comparisons of wind-tunnel and flight-measured values of stability and control characteristics are of considerable interest to the designer, since the wind-tunnel method of testing is one of the prime sources upon which estimates of the characteristics of a new configuration are based. In this paper comparisons are made of some of the more important stability and control characteristics of three swept-wing airplanes as measured in flight and in wind tunnels. Wind-tunnel data are used from high-speed closed-throat tunnels, a slotted-throat transonic tunnel, and a supersonic tunnel. The comparison shows that, generally speaking, the wind tunnels predict all trends of characteristics reasonably well. There are, however, differences in exact values of parameters, which could be attributed somewhat to differences in the model caused by the method of support. The small size of the models may have some effect on measurements of flap effectiveness. When non-linearities in derivatives occur during wind-tunnel tests, additional data should be obtained in the region of the non-linearities. Also, non-linearities in static derivatives must be analyzed on the basis of dynamic motions of the airplane. Aeroelastic corrections must be made to the wind-tunnel data for models of airplanes which have thin surfaces and are to be flown at high dynamic pressures. Inlet effects can exert an influence on the characteristics, depending upon air requirements of the engine and location of the inlets.
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    Report Date: August 1956
    No. Pages: 20


  2. SKIN AND STRUCTRUAL TEMPERATURE MEASUREMENTS ON RESEARCH AIRPLANES AT SUPERSONIC SPEEDS
    Authors: Richard D. Banner and Frank S. Malvestuto, Jr.
    Report Number: NACA-H56BANN
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NACA High-Speed Flight Station is presently engaged in a comprehensive program to investigate and analyze the aerodynamic heating of airplanes in flight at supersonic speeds. Skin and structural temperatures have been determined in flight by use of thermocouples and temperature resistance gages installed on various research airplanes. Such data have recently been obtained on two research airplanes, the Bell X-2 and the Bell X-1B. The object of this paper is to show some of the actual magnitudes and trends in the structural temperatures that exist in an airplane experiencing the effects of aerodynamic heating. Although time has not permitted an analysis of much of these temperature data, the information presented herein is a cross section of the work that is presently being done. The data presented do not cover the speed and altitude range anticipated for the proposed North American X-15; however, they provide actual full-scale, experimental, structural temperature information for comparison with present analytical and wind-tunnel studies. Shown in figure 1 is a sketch of the two airplanes. The X-2 airplane is constructed primarily of steel. Nose-cone skin temperatures have been obtained on this airplane at supersonic speeds up to a Mach number of 3.2. The X-1B is constructed primarily of aluminum. The maximum speed capability of the X-1B is less than that of the X-2. Very extensive temperature measurements are being made throughout the structure of the airplane. To date, data have been obtained in flight to a Mach number of 1.8 at about 300 locations throughout the airplane.
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    Report Date: October 1956
    No. Pages: 13
    Keywords:      X-15 aircraft; Research and development; Research aircraft; Aerodynamic heating; Supersonic speed; Aircraft construction materials; Steels; Aluminum alloys; Aircraft structures; Temperature; Skin temperature (non-biological); Temperature distribution


  3. FLIGHT EXPERIENCE WITH PRESENT RESEARCH AIRPLANES
    Authors: Hubert M. Drake
    Report Number: NACA-H56DRAK
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In summary, it has been shown that several of the problems of direct pertinence to the X-15 project have been experienced on current research airplanes. The future investigations of the handling qualities of the X-1B and X-1E will furnish additional information. The experiences with the X-1A and the X-2 airplanes are indicative of the extreme caution that is required in this type of flight research. Critical conditions with the X-1B and X-1E will be approached with great care. The X-1B is to be used in investigations of handling qualities at high altitudes and low dynamic pressure. The flights with this airplane will probably not involve Mach numbers much above 2 because of the loss of directional stability. The investigation of control at very low dynamic pressure will be extended to include rocket reaction controls. Initially, the X-1E program will be to investigate the stability and control characteristics in the Mach number range above 2 and will include means of improving directional stability and handling at high angles of attack. At a later date, the X-1E wll be used to extend the low-dynamic-pressure investigation of the X-1A.
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    Report Date: October 1956
    No. Pages: 9
    Keywords:      X-15 aircraft; Research aircraft; Flight tests; Aircraft performance; Research and development; Thrust; Misalignment; Dynamic stability; Directional stability; Angle of attack; Controllability; High altitude; X-2 aircraft


  4. GENERAL BACKGROUND OF THE X-15 RESEARCH-AIRPLANE PROJECT
    Authors: Dr. Hugh L. Dryden
    Report Number: NACA-H56DRYD
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Paper presents the general background that led to the establishment of a committee and their subsequent recommendation that a research aircraft, X-15 aircraft, be constructed to meet the needs of maintaining supremacy in the air.
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    Report Date: October 1956
    No. Pages: 6
    Keywords:      Conferences; Histories; Research aircraft; Research and development; X-15 aircraft


  5. STUDIES OF REACTION CONTROLS
    Authors: Wendell H. Stillwell
    Report Number: NACA-H56STIL
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The attitude-control method selected for the North American X-15 for flight at extremely low and zero dynamic pressures utilizes the reaction forces developed by small-rocket units located on the airplane to produce rolling, pitching, and yawing mements. An investigation of reaction control similar to those selected for the X-15 has shown that unique control problems exist for flight at the low dynamic pressures where this type of control is used. Although the Bell X-15 configuration was utilized for this investigation, a range of variables was covered to determine the significant effects of various factors on flight with reation controls. It was also of interest to determine fuel requirements for the rocket units. The investigation consisted of analog-computer studies and ground-simulator tests. The significant results of this investigation is discussed in this paper.
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    Report declassified: 25 April 1967
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    Report Date: October 1956
    No. Pages: 8
    Keywords:      X-15 aircraft; Research and development; Research aircraft; Aircraft control; Attitude control; Reaction control; Analog simulation; Flight simulators; Control simulation. Lateral control;Longitudinal control; Yaw; Low pressure; Dynamic Pressure


  6. INSTRUMENTATION FOR THE X-15
    Authors: I. Taback and G.M. Truszynski
    Report Number: NACA-H56TABA
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The development of a research airplane which extends manned flight into regions where extremes of temperature and pressure are reached requires the simultaneous development of new instrumentation technique not only to insure safe operation of the aircraft but also to derive a maximum of research data throughout the operational range of the aircraft. The instrumentation required for the North American X-15 airplane project consists of ground range and its associated equipment and airborne equipment required for pilot's displays and for research measurements. This paper outlines a plan for a ground range, which is based upon developed equipment already in use, and also discusses the airborne instrumentation and some of the special airborne devices which are made necessary by the extended performance capabilities of the airplane.
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    Report Date: October 1956
    No. Pages: 11
    Keywords:      X-15 aircraft; Research and development; Research aircraft; Aaircraft instruments; Ground support equipment; Airborne equipment; Avionics; measuring instruments; Ranges


  7. AN ANALOG STUDY OF THE RELATIVE IMPORTANCE OF VARIOUS FACTORS AFFECTING ROLL COUPLING
    Authors: Joseph Weil and Richard E. Day
    Report Number: NACA-RM-H56A06
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An analog study of the roll-coupling problem has been made for a representative swept-wing and tailless delta-wing configuration. The investigation, made primarily for subsonic flight conditions, included the determination of the effects of wide variations in many of the pertinent aerodynamic derivatives on the motions developed in rolling maneuvers. The influence of large changes in principal axis inclination and mass distribution is also considered. The results indicated that as first predicted in NACA TN 1627 the relationship between the longitudinal and directional stability is of paramount importance. For most current designs an optimum condition exists when the natural frequencies in pitch and yaw are approximately equal. Increases in pitch damping had a pronounced favorable effect in reducing the amplitudes of the motions encountered and were, in general, considerably more effective than corresponding increases in yaw damping. practical redistribution of mass produced only relatively monor changes in the overall results. It was found that the amplitude of the motions developed for a given aileron deflection depends to a large extent on the duration of the maneuver (change in bank angle). Limited studies indicated that 90 degree rolls. The angle of attack of the principal axis has an important bearing on the behavior, particularly in the absence of other disturbing functions. If the initial angle of attack is maintained constant, a reduction in altitude will delay critical conditions to a higher roll rate but the maximum amplitudes may be only slightly affected. Small inadvertent stablilzer inputs can greatly affect the agnitude of the motions developed. It is difficult to geralize on the effects of Mach number variation because this variable affects many of the controlling parameters. Utilizing simple concepts proved useful in assessing the qualitative effects of many of the aerodynamic and inertia pparameters and changes in flight condition. The calculated lower resonant frequency generally corresponded to the average roll velocity at which the more serious motions could be expedted. It is indicated that rational design procedure can avoid the problem of serious roll coupling at supersonic speeds and minimize the problem in the subsonic speed range for theconfigurations of the type cosidered.
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    Report Date: April 1956
    No. Pages: 82


  8. FLIGHT MEASUREMENTS OF HORIZONTAL-TAIL LOADS ON THE DOUGLAS X-3 RESEARCH AIRPLANE
    Authors: Harriet J. Stephenson
    Report Number: NACA-RM-H56A23
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight measurements of the horizontal-tail loads on the Douglas X-3 reserch airplane during wind-up turns, pull-ups, and stabilizer pulses were made over an altitude range from 27,000 to 33,000 feet and throughout a Mach number range from 0.65 to 1.16. The results of these measurements are presented in this paper. The normal-force-curve slope of the horizontal-tail panel (CNa)t' derived from stabilizer pulses, had a maximum value of 0.082 and occurred at a Mach number of 0.925. At a Mach number of 1.00 the value of the slope decreased to 0.063 and for higher Mach numbers again increased with Mach number. Balancing-tail loads, downwash at the tail, and total airplane pitching monents were obtained from pull-ups and wind-up turns. Balancing-tail loads varied nonlinearly with airplane normal-force coefficient throughout the lift range; the wing fuselage ws stable for the moderate lift range with increasing stability for increasing Mach number. An increase in stability occurred at lift coefficients between 0.2 and 0.4. The wing-fuselage became unstable at the high lift coefficients. Down wash varied nonlinearly with angle of attack. An increase in the variation of downwash with angle of attack de/da or a decrease in tail stability occurred at angles of attack between 4 degrees and 8 degrees. The total airplane pitching moment also displayed nonlinear variations with angle of attack. The airplane became unstable at angles of attack between 7 degrees and 13 degrees.
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    Report Date: April 1956
    No. Pages: 34


  9. TIME-VECTOR DETERMINED LATERAL DERIVATIVES OF A SWEPT-WING FIGHTER-TYPE AIRPLANE WITH THREE DIFFERENT VERTICAL TAILS AT MACH NUMBERS BETWEEN 0.70 AND1.48
    Authors: Cheste H. Wolowicz
    Report Number: NACA-RM-H56C20
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: As part of the flight research program conducted on a swept-wing fighter-type airplane, rudder-pulse maneuvers were performed at altitudes from 30,000 to 43,000 feet over a Mach number range of 0.71 to 1.48 to determine the lateral stability characteristics relative to the stability axes, in general, and the lateral derivative characteristics, in particular. The time-vector method of analysis was used. Four configurations were employed in the investigation. Three configurations involved three different vertical tails with varying aspect ratio or area, or both. The fourth configuration employed a large tail, which had been used in the third configuration, and an extension of the wing tips.
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    Report declassified: 22 July 1959
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    Report Date: June 1956
    No. Pages: 95


  10. FLIGHT DETERMINATION OF THE LATERAL HANDLING QUALITIES OF THE BELL X-5 RESEARCH AIRPLANE AT 58.7 DEGREES SWEEPBACK
    Authors: Thomas W. Finch and Joseph A. Walker
    Report Number: NACA-RM-H56C29
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The Bell X-5 variable-sweep research airplane has been tested primarily at 58.7 degrees sweepback to determine the characteristics at transonic speeds of a fighter-type airplane having extreme sweepback. Some of the dynamic and static lateral stability characteristics have been discussed previously. This paper will summarize the overall lateral stability and control characteristics up to a Mach number of 0.97 at 40,000 feet and to slightly lower Mach numbers at altitudes of 25,000 and 15,000 feet. The dynamic characteristics were influenced by aerodynamic and engine gyroscopic coupling. The short-period lateral oscillations were moderately well damped up to a Mach number of 0.80, but were only tolerable at higher Mach numbers because of the influence of nonlinear damping. However, the damping was generally unsatisfactory over most of the Mach number range when compared to the Military Specification. The apparent directional stability was positive and about constant for all test altitudes up to a Mach number of 0.85 and increased appreciably at higher Mach numbers. The apparent effective dihedral was positive and had a high value, increasing rapidly at higher Mach numbers. The lateral-force coefficient per degree of sideslip was about constant for all altitudes to a Mach number of 0.94 and increased rapidly with further increase in Mach number at 40,000 feet. There was little change in pitching moment caused by sideslip at any altitude for the limited range of sideslip angles tested. Changes in dynamic pressure had little effect on most of the static stability characteristics. The rolling characteristics were affected considerably by the adverse dihedral effects at some flight conditions. The aileron effectiveness was low at all altitudes and varied little with Mach number. The airplane failed to meet the Military Specification requirement for rolling velocity and the requirement of 1 second to bank to 100 degrees.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 28 October 1960
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    Report Date: May 1956
    No. Pages: 42
    Keywords:      X-5 aircraft


  11. EFFECT OF SEVERAL WING MODIFICATIONS ON THE SUBSONIC AND TRANSONIC LONGITUDINAL HANDLING QUALITIES OF THE DOUGLAS D-558-II RESEARCH AIRPLANE
    Authors: Jack Fischel and Donald Reisert
    Report Number: NACA-RM-H56C30
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The subsonic and transonic longitudinal handling qualities of the Douglas D-558-II research Airplane were measured with several wing modifications designed to alleviate swept-wing instability and pitch-up. The airplane conigurations investigated included the basic wing configuration and two wing-fence configurations in combination with retracted, free-floating, or extended stats, and a wing leading-edge chord-extension configuration. All configurations were tested in the clean condition. None of the wing modifications had an appreciable effect on the decay in stick-fixed stability (pitch-up) exhibited by the airplane at moderate angles of attack, and all configurations were considered by the pilots to be unsatisfactory and uncontrollable in the pitch-up region. Both flight and wind-tunnel results indicated that the position of the horizontal tail should be lowered appreciably to obtain substantial improvement in longitudinal handling qualities of the airplane. Wing fences had no apparent effect on airplane buffeting characteistics with slats retracted. With wing slats free to float, the onset of buffeting was delayed at low Mach numbers, whereas buffeting was generally seriously aggravated by wing chord-extensions. Fully extending the wing slats had no appreciable effect on buffeting at low and moderate lifts but delayed the intensity rise to higher lift levels.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 24 June 1958
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    Report Date: June 1956
    No. Pages: 60
    Keywords:      Douglas D-558-II aircraft


  12. LIFT AND DRAG OF THE BELL X-5 RESEARCH AIRPLANE IN THE 45 DEGREE SWEPTBACK CONFIGURATION AT TRANSONIC SPEEDS
    Authors: Jack Nugent
    Report Number: NACA-RM-H56E02
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight investigation was made of the Bell X-5 variable-sweep research airplane in the 45 degree wing sweptback configuration over a Mach number range from 0.61 to 1.01. Lift and drag values are presented and a comparison is made with data previously obtained with a wing sweepback of 59 degrees. For the configuration at a wing sweep of 45 degrees the lift-curve slope remained constant at a value of about 0.067 degrees-1 as Mach number increased from 0.61 to 0.80 and increased to value of 0.078 as Mach number increased further to 0.95. Over the Mach number range tested, the configuration at 45 degrees sweepback had a lift-curve slope approximately 0.03 degree-1 higher than the slope for the 59 degree configuration. Below the drag rise the 45 degree configuration had a zero-lift drag coefficient of 0.020 as compared with a zero-lift drag coefficient of 0.0175 for the 59 degree sweptback configuration. The drag-rise Mach number was 0.85 for the configuration at 45 dgrees sweepback as compared to 0.90 for the wing at 59 degrees sweepback. At 45 degrees wing sweep the lift-drag ratio exceeded that for the 59 degrees sweep for a Mach number range from 0.61 to 0.88 with a maximum difference of about 0.7 at a Mach number of 0.82, but was less for Mach numbers in excess of 0.88. The drag-due-to-lift factor for the 45 degree sweptback configuration was constant at approximately 0.18 as Mach number increased from 0.61 to 0.94. This value of drag-due-to-lift factor was about 0.12 less than that for the 59 degree sweptback configuration.
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    Report declassified: 1 September 1959
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    Report Date: July 1956
    No. Pages: 24
    Keywords:      X-5


  13. FLIGHT-DETERMINED TRANSONIC LIFT AND DRAG CHARACTEISTICS OF THE YF-102 AIRPLANE WITH TWO WING CONFIGURATIONS
    Authors: Edwin J. Saltzman, Donald R. Bellman and Norman T. Musialowski
    Report Number: NACA-RM-H56E08
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Lift and drag characteristics of the Convair YF-102 airplane have been determined in flight for the symmetrical wing configuration and for the cambered wing configuration. The data were obtained for lift coefficients between 0.025 and 0.73, for altitudes of 25,000 feet, 40,000 feet, and 50,000 feet and for Mach numbers from 0.6 to 1.17. The results indicated that the lift-curve slopes increased gradually with lift over the lift range from 0.1 to 0.4 with much greater increase for the symmetrical wing configuration than for the cambered wing configuration. In addition, the modifications comprising the cambered configuration caused the angle of attack for zero lift to increase less than 0.5 degrees. The cambered configuration experienced lower drag coefficient values for lift coefficient values above 0.1. Maximum advantage of the cambered configuration was realized at lift coefficients of 0.3 and above, where the reduction in drag coefficient amounted to about 0.01. The drag-due-to-lift values for the cambered configuration were 65 to 75 percent of the symmetrical values at a lift coefficient of 0.2 and for Mach number values below the drag rise. At a lift coefficient value of 0.35 the drag-due-to-lift of the cambered wing was 75 to 85 percent of the symmetrical wing values. The maximum lift-drag ratio for the cambered wing was almost 20 pecent higher than for the comparable symmetrical wing values throughout the Mach number range. Comparisons of flight and tunnel drag characteristics suggest a tendency of zero-lift drag coefficient to decrease with increasing Reynolds number; however, it cannot be determined from these comparisons what part of the zero-lift drag coefficient change is a result of Reynolds number and what portion should be attributed to model variations from exact reproductions, or to inaccuracies in the data.
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    Report declassified: 12 January 1961
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    Report Date: July 1956
    No. Pages: 45
    Keywords:      YF-102 aircraft


  14. CORRELATION OF FLIGHT AND ANALOG INVESTIGATIONS OF ROLL COUPLING
    Authors: Joseph Weil and Richard E. Day
    Report Number: NACA-RM-H56F08
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A brief review of NACA flight experience relating to the roll-coupling problem is presented. Conditions rated by pilots as intolerable, marginal, and good are discussed and correlated with calculated results. A suggested flight test procedure for roll-coupling investigatons and a discussion of several other items of general interest are also presented. Good correlation was obtained between calculated motions and flight data in a number of instances. It would appear that intolerable conditions should be predictable from general analog studies. The primary difference between the marginally acceptable and intolerable roll-coupled maneuvers would appear to be the much larger negative normal acceleration attained in the latter maneuvers, as well as a somewhat higher sideslip angle. The suggested approach of close coordination of flight test results with calculations should greatly lessen the possibilty of encountering an upredictable violent roll-coupled maneuver.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 28 July 1960
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    Report Date: September 1956
    No. Pages: 20
    Keywords:      Roll coupling


  15. FLIGHT EVALUATION OF THE LATERAL STABILITY AND CONTROL CHARACTERISTICS OF THE CONVAIR YF-102 AIRPLANE
    Authors: Thomas R. Sisk, William H. Andrews and Robert W. Darville
    Report Number: NACA-RM-H56G11
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The lateral stability and control characteristics were investigated on the Convair YF-102 airplane during flights of the National Advisory Committee for Aeronautics research program. The investigation included gradually increasing sideslips, rudder-fixed aileron rolls, rudder pulses, and trim runs at altitudes of 25,000 and 40,000 feet over the test Mach number range. A few wind-up turns were performed at an altitude of 50,000 feet to investigate directional stability at high lifts. The lateral handling characteristics appeared satisfactory when viewed in terms of gradually increasing sideslips. A large directional trim change was encountered at all altitudes at a Mach numer of approximately 0.95 and a directional divergence was encountered at high lifts(angle of attack approximately 20 degrees). The lateral dynamic stability characteristics were generally unsatisfactory but more tolerable at the higher speeds. Violent inertial coupling was encountered during aileron rolls at a Mach number of 0.74; however, no difficulty was encountered when observing the restriction of rate of roll of 100 degrees per second and angle of bank of 100 degrees.
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    Report Date: January 1956
    No. Pages: 39
    Keywords:      YF-102 aircraft


  16. WING LOADS AND LOAD DISTRIBUTIONS THROUGHOUT THE LIFT RANGE OF THE DOUGLAS X-3 RESEARCH AIRPLANE AT TRANSONIC SPEEDS
    Authors: Earl R. Keener and Gareth H. Jordan
    Report Number: NACA-RM-H56G13
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Wing loads and load distributions were obained by differential-pressure measurements between the upper and lower surfaces of the left wing of the Douglas X-3 research airplane to derermine the effects of angle of attack and Mach number on the wing characteristics at transonic Mach numbers. The wing has an aspect ratio of 3.09 and a modified 4.5-percent-thick hexagonal section. Data cover the range from near-zero lift to maximum lift and from a Mach number of 0.71 to a Mach number of 1.15. The chordwise load distributions and the wing-section aerodynamic characteristics were similar at each wing station. A large load developed at the leading edge resulting from the relatively sharp leading edge. At Mach numbers below 0.9 separation of the flow from the leading edge resulted in a loss in leading-edge load and a low maximum lift. The maximum normal-force coefficient of the wing panel was 0.66 at a Mach number of 0.71 compared to 1.2 at supersonic Mach numbers. Spanwise load distributions were essentially elliptical throughout the lift and Mach number range tested. Values of normal-force-curve slope range from 0.076 per degree at a Mach number of 0.71 to 0.116 per degree at a Mach number of 1.0. Variation of pitching moment with lift was unstable at the lower Mach numbers, becoming increasingly stable above a Mach number of about 0.9. The chordwise location of the center of pressure varied with angle of attack between 15- and 30-pecent chord at subsonic Mach numbers and between 31- and 37-percent chord at supersonic Mach numbers. The spanwise location of the center of pressure was relatively constant with lift and Mach number at about 42 percent of the panel span. The flight results are in good agreement with wind-tunnel results at Mach numbers below 0.90 and in fair agreement at Mach numbers of 0.90 and 0.92. Deflecting the leading-edge flap about 7 degrees over a Mach number range of 0.71 to 0.80 increased the maximum normal-force coefficient about 0.06 and moved the center of pressure rearward at the lower angles of attack and slightly forward at the higher angles of attack. No change occurred in the spanwise location of the center of pressure.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 17 July 1958
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    Report Date: November 1956
    No. Pages: 192
    Keywords:      X-3 aircraft


  17. DYNAMIC LONGITUDINAL STABILITY CHARACTERISTICS OF A SWEPT-WING FIGHTER-TYPE AIRPLANE AT MACH NUMBERS BETWEEN 0.36 AND 1.45
    Authors: Chester H. Wolowicz
    Report Number: NACA-RM-H56H03
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: As part of the fight research program conducted by the National Advisory Committee for Aeronautics on a swept-wing fighter-type airplane not equipped with an automatic pitch damper, pulse maneuvers were performed at altitudes from 10,000 to 40,000 feet over a Mach number range from 0.36 to 1.45 to determine the longitudinal stability characteristics and derivatives for an original-wing and an extended wing-tip configuration. The longitudinal dynamic behavior of the airplane during simulated combat maneuvers at altitudes of 30,000 to 40,000 feet was not considered satisfactory, especially at supersonic speeds, becuase of insufficient pitch damping. The addition of the wing-tip extensions caused a slight favorable shift in the aerodynamic center of the airplane. The static margin of the extended wing-tip configuration is of the order of 12-percent mean aerodynamic chord in the subsonic region and 29-percent mean aerodynamic chord at Mach numbers above 1.2. Wind-tunnel data for the two wing configurations investigated showed good agreement with transonic flight results for the lift-curve slope and the static stability derivative C (sub m, sub sub alpha); poor agreement was evident in the supersonic regon.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 26 July 1957
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    Report Date: April 1956
    No. Pages: 29


  18. ANALYSIS OF THE VERTICAL-TAIL LOADS MEASURED DURING A FLIGHT INVESTIGATION AT TRANSONIC SPEEDS OF THE DOUGLAS X-3 RESEARCH AIRLANE
    Authors: William L. Marcy, Harriet J. Stephenson and Thomas V. Cooney
    Report Number: NACA-RM-H56H08
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Results are presented of an analysis of the strain-gage measurements of vertical-tail loads experienced in rudder pulses, gradually increasing sideslips, and rudder-fixed aileron rolls at transonic speeds with the Douglas X-3 research airplane. A Mach number range from approximately 0.7 to 1.2 at an altitude of about 30,000 feet was covered during this investigation. The lift-curve slope of the vertical tail increased with increasing Mach number from a value of 0.038 per degree at a Mach number of 0.7 to a maximum of 0.048 per degree at a Mach number of 0.94, followed by a reduction to 0.041 at supersonic Mach numbers. A comparison with available methods of estimating this parameter indicated good agreement with flight results. The effectiveness of the rudder (lift-curve slope of the vertical tail due to deflecting the rudder) decreased from approximately 0.020 per degree at subsonic Mach numbers to 0.013 per degree at supersonic Mach numbers. In several violent roll maneuvers, sideslip angles of as much as 21 degrees were reached without stalling of the vertical tail, although vertical-tail effectiveness was reduced at sideslip angles above about 12 degrees. At sideslip angles below about 6 degrees the center of pressure of the load on the vertical tail in sideslip was practically unchanged with Mach number, remaining at about 55 percent span and 30 percent chord. At sideslip angles above this value a more rearward center of pressure was indicated. The center of load due to displacing the rudder was farther inboard, 45 percent span, and farther rearward, 63 percent chord, than in sideslip, and moved rearfward to 85 percent chord at supersonic Mach numbers. The variation of airplane yawing-moment coefficient with sideslip, as determined from vertical-tail-loads measurements, increased from 0.0023 per degree at a Mach numbr of 0.7 to a maximym of 0.0032 at a Mach number of 0.94, and decreased to 0.0023 at supersonic Mach numbers.
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    Report Date: November 1956
    No. Pages: 33
    Keywords:      X-3 aircraft; Vertical-Tail loads; Transonic speeds


  19. LONGITUDINAL STABILITY CHARACTERISTICS OF THE CONVAIR YF-102 AIRPLANE DETERNINED FROM FLIGHT TESTS
    Authors: William H. Andrews, Thomas R. Sisk and Robert W. Darville
    Report Number: NACA-RM-H56I17
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An analysis was made of the longitudinal stability characteristics of the cambered-wing version of the Convair YF-102 airplane from flight data obtained up to a Mach number of 1.18 at altitudes of 25,000, 40,000, and 50,000 feet. In addition, trim data are analysed for the symmetrical-wing configuration at the two lower altitudes. The longitudinal control for trim appears conventional, with the unstable region occurring generally in the Mach number rnge from 0.87 to 0.95. The cambered-wing modification reduced the elevator required for 1 g trim below that required with the original-wing configuration by approximately 0.6 degrees to 1.9 degrees at 25,000 feet. The longitudinal damping characteristics met the Military Specification to damp to one-half amplitude in 1 cycle, but did not indicate that damping to one-tenth amplitude in 1 cycle could be attained. The pilots commented that the damping was insufficient. Generally there was a gradual decrease in stability with increasing lift. However, no severe pitch-up tendencies were exhibited, except when accelerating or decelerating through the trim-change region. The stability more than doubles between Mach numbers of 0.60 and 1.16; however, the control effectiveness shows an increase up to a Mach number of 0.89 with a rapid decrease of approximately 50 pecent occurring between Mach numbers of 0.90 and 1.0. An abrupt decrease in the stick-free stability exhibited between 1.5g and 2.0g is felt to result from the location of the total head and static-sensing probes for the Mach compensating instrument of the artificial-feel system.
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    Report declassified: 28 July 1960
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    Report Date: January 1956
    No. Pages: 41
    Keywords:      YF-102 aircraft


  20. AN ANALYSIS OF ESTIMATED AND EXPERIMENTAL TRANSONIC DOWNWASH CHARACTERISTICS AS AFFECTED BY PLAN FORM AND THICKNESS FOR WING AND WING-FUSELAGE CONFIGURATIONS.
    Authors: Joseph Weil, George S. Campbell and Margaret S. Diederich
    Report Number: NACA-TN-3628
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper presents a summary of the effects of changes in wing plan form and thickness ratio on the downwash characteristics of wing and wing-fuselage configurations in the Mach number range between 0.6 and 1.1. Data obtained by the transonic-bump technique at two tail heights have been compared with theoretical estimations made in the subsonic and supersonic Mach number range.
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    Report Date: April 1956


  21. FLIGHT TECHNIQUES FOR DETERMINING AIRPLANE DRAG AT HIGH MACH NUMBERS
    Authors: De E. Beeler, Donald R. Bellmen and Edwin J. Saltzman
    Report Number: NACA-TN-3821
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The measurement of total airplane drag in flight is necessary to assess the applicability of wind-tunnel model data. The NACA High-Speed Flight Station has investigated and developed techniques for measuring the drag of high-speed research aireplanes and current fighter-type airplanes. The accelerometer method for detemining drag was found to be the most satisfactory method for research work, because it is the only method permitting a complete coverage of the Mach number and angle-of-attack capabilities of and airplane. Deterniming drag by the accelerometer method requires the accurate measurement of longitudinal and normal accelerations, angle of attack, and engine thrust. In addition, the static presure, airspeed, airplane weight, and longitudinal control positions must be measured. The accurate measurement of longitudinal and normal acceleration can be make and recorded by means of specially constructed mechanical accelerometers that have been developed by the NACA. Fuselage nose booms are used to reduced the flow-field errors in the measurement of static pressure, airspeed, and angle of attack. The errors can be reduced further to an acceptable level by established calibration techniques. Satisfactory methods for determining in flight the thrust of turbojet-afterburner and rocket engines are available. The flight drag data generally can be separated into components consisting of trim, skin-friction, pressure-induced, and wave drags. The comparison of flight and wind-tunnel data must be made on the basis of component drags it a proper interpretation of the results is to be obtained.
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    Report Date: August 1956
    No. Pages: 41


  22. COMPARISON OF FLIGHT AND WIND-TUNNEL MEASUREMENTS OF HIGH-SPEED-AIRPLANE STABILITY AND CONTROL CHARACTERISTICS
    Authors: Walte C. Williams, Hubert M. Drake and Jack Fischel
    Report Number: NACA-TN-3859
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Comparisons of wind-tunnel and flight-measured values of stability and control characteistics are of considerable interest to the designer, since the wind-tunnel method of testing is one of the prime sources upon which estimates of the characteristics of a new configuration are based. In this paper comparisons are made of some of the more important stability and control characteristics of three swept-wing airplanes as measured in flight and in wind tunnels. Wind-tunnel data from high-speed closed-throat tunnels, a slotted-throat transonic tunnel, and a supersonic tunnel are used. The comparisons show that, generally speaking, the wind tunnels predict all trends of characteristics reasonably well. There are, however, differences in exact values of parameters, which could be attributed somewhat to differences in the model caused by the method of support. The small size of the models may have some effect on measurements of flap effectiveness. When nonlinearities in derivatives occur during wind-tunnel tests, additional data should be obtained in the region of the nonlinearities in order to predict more accurately the flight characteristics. Also, nonlinearities in static derivatives must be analyzed on the basis of dynamic motins of the airplane. Aeroelastic corrections must be made to wind-tunel data for models of airplanes which have thin surfaces and are to be flown at high dynamic pressures. Inlet effects can exert an influence on the characteristics, depending upon air requiements of the engine and location of the inlets.
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    Report Date: August 1956
    No. Pages: 17
 
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