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  1. AEROELASTIC EFFECTS OF AERODYNAMIC HEATING
    Authors: Hugh L. Dryden and Dr. John E. Duberg
    Report Number: AGARD20-P10
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In conclusion, it appears that the design of aircraft to withstand aeroelastic difficulties at high supersonic speeds will of necessity require the consideration of the effect of aerodynamic heating. Among the various aeroelastic consequences of aerodynamic heating, the reduction of over-all stiffness through the action of thermal stress is the most novel and may well turn out to be the most serious. An appreciation of this phenomenon must become part of the working equipment of the modern aeroelastician.
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    Report Date: June 1955
    No. Pages: 10
    Keywords:      Aerodynamic heating; Aeroelasticity; Structural reliability; Aircraft design; Airframe materials; Conferences; Flutter analysis; Supersonic speed


  2. FLIGHT EXPERIENCE OF INERTIA COUPLING IN ROLLING MANEUVERS
    Authors: Joseph Weil, Ordway B. Gates, Jr., Richard D. Banner and Albert E. Kuhl
    Report Number: NACA-H55WEIL
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Violent coupled lateral-longitudinal motions have been encountered in flight on two airplanes during abrupt aileron rolls at relatively high speed. During these motions, various structured design loads and load factors were either exceeded or approached. It was demonstrated on one airplane that the motions can be approximated reasonably well by using a five-degree-of-freedom analysis. From flight tests of the swept-wing airplane at relatively high altitude, it was found that the severity of the divergent tendency increased with roll velocity and was sensitive to roll direction and stabilizer input. Calculated results indicated that considerably more critical conditions from the loads standpoint can be expected at lower altitudes when the roll is initiated from a pull-up condition. Perhaps one of the fundamental reasons for the occurrence of the large motions on both airplanes was the presence of insufficient directional stability. Doubling the directional stability level of the swept-wing airplane resulted in substantially improved flight characteristics; but calculations indicated that, if the tail size is increased beyond a certain point, considerably higher tail loads and larger peak normal accelerations can be obtained than with a tail affording a somewhat lower level of stability. At present, analytical investigations are under way to enable a better understanding of the overall problem of coupled lateral-longitudinal motions in rolling maneuvers. It is not yet known whether a practical design approach exists that would produce desirable characteristics for a large range of flight conditions without the sacrifice of performance or the resort to artificial stabilization. It is also true that coupling can have a large effect on the predicted loads, even for configurations that have satisfactory handling qualities; therfore, the coupling of the lateral and longitudinal degrees of freedom should be considered for load evaluations of rolling maneuvers on most high-speed airplanes.
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    Report Date: March 1955
    No. Pages: 12


  3. ROLLING PERFORMANCE OF THE REPUBLIC YF-84F AIRPLANE AS MEASURED IN FLIGHT , Research Memorandum
    Authors: John B. McKay
    Report Number: NACA-RM-H54G20A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight measurements of the rolling performance of the Republic YF-84F airplane were made at altitudes of 10,000, 25,000, and 40,000 feet. The tests were conducted over a Mach number range from 0.35 to 0.95.
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    Report Date: January 1955


  4. FLIGHT MEASUREMENTS OF ELEVON HINGE MOMENTS ON THE XF-92A DELTA-WING AIRPLANE , Research Memorandum
    Authors: Clinton T. Johnson and Albert E. Kuhl
    Report Number: NACA-RM-H54J25A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Elevon hinge-moment measurements were made during flight tests of the Convair XF-92A delta-wing airplane over the Mach number range from 0.70 to 0.95. Hinge moments were measured during longitudinal elevon pulses, aileron rolls, and wind-up turns. Data are presented giving the variation of C (sub) h (sub) delta and C (sub) h (sub) alpha as determined from these tests.
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    Report Date: January 1955


  5. FLIGHT MEASUREMENTS OF THE DYNAMIC LONGITUDINAL STABILITY AND FREQUENCY-RESPONSE CHARACTERISTICS OF THE XF-92A DELTA-WING AIRPLANE , Research Memorandum
    Authors: Euclid C. Holleman and William C. Triplett
    Report Number: NACA-RM-H54J26A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Results of dynamic longitudinal flight test conducted with the XF-92A delta-wing airplane over a Mach number range of 0.42 to 0.94 at an altitude of about 30,000 feet are presented. The data were analyzed by measuring the airplane oscillatory characteristics, by matching the airplane system with an analog computer, and by determining the frequency-response characteristics of the airplane. Wherever possible, stability derivatives were computed and are presented as a function of Mach number.
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    Report Date: January 1955


  6. LONGITUDINAL STABILITY CHARACTERISTICS IN MANEUVERING FLIGHT OF THE CONVAIR XF-92A DELTA-WING AIRPLANE INCLUDING THE EFFECTS OF WING FENCES , Research Memorandum
    Authors: Thomas R. Sisk and Duane O. Muhleman
    Report Number: NACA-RM-H54J27
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The longitudinal maneuvering stability characteristics are evaluated on the Convair F-92A delta-wing airplane in wind-up turns over the Mach number range from 0.70 to 0.95 at altitudes between 22,000 and 39,000 feet. A longitudinal stability reduction evidenced as a pitch-up encountered over the Mach number range tested is evaluated along with the airplane behavior in the region of reduced stability. Two wing fence configurations are evaluated and compared with the basic airplane configuration characteristics.
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    Report Date: January 1955


  7. THE EFFECT OF THE BLUNT-TRAILING-EDGE ELEVONS ON THE LONGITUDINAL AND LATERAL HANDLING QUALITIES OF THE X-4 SEMITAILLESS AIRPLANE , Research Memorandum
    Authors: Edwin J. Saltzman
    Report Number: NACA-RM-H54K03
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The effects of thickening the trailing edges of the elevons on the static longitudinal stability and control, lateral control for the X-4, a swept-wing semitailless airplane, are presented. The results of this study are compared with similar tests for the X-4 with conventional elevon trailing edges.
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    Report Date: January 1955


  8. RESULTS OF MEASUREMENTS MADE DURING THE APPROACH AND LANDING OF SEVEN HIGH-SPEED RESEARCH AIRPLANES
    Authors: Wendell H. Stillwell
    Report Number: NACA-RM-H54K24
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An investigation has been conducted by the National Advisory Committee for Aeronautics of the landing characteristics of the X-1, X-3, and D-558-I straight-wing, the X-4, X-5, and D-558-II swept-wing, and the XF-92A delta-wing high-speed research airplanes. These tests have shown that ground contact occurs at about 70 to 90 percent of the maximum normal-force coefficient even though the maximum normal-force coefficient was established by maximum fligt, stability or control limitations, or ground clearance restrictions. The average vertical velocity at ground contact for the normal landings was about 2 feet per second and the maximum vertical velocity was about 4.6 feet per second. Tests of the X-4 airplane to determine the effect of lift-drag ratio on the landing maneuver showed that the largest portion of the landing flare was made at altitudes above 50 feet at low lift-drag ratios and that, although the vertical velocities during the approach varied from 30 to 90 feet per second, the vertical velocities at contact were less than 5.5 feet per second.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 17 July 1958
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    Report Date: February 1955
    No. Pages: 25


  9. FLIGHT MEASUREMENTS AT TRANSONIC SPEEDS OF THE BUFFETING CHARACTERISTICS OF THE XF-92A DELTA-WING RESEARCH AIRPLANE , Research Memorandum
    Authors: Thomas F. Baker and Wallace E. Johnson
    Report Number: NACA-RM-H54L03
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Measurements were made on the XF-92A delta-wing airplane of buffet-induced fluctuations in normal acceleration at the airplane center of gravity and of fluctuations in wing structural shear load in the Mach number range from 0.6 to 0.96 at altitudes from 25,000 to 38,000 feet. Airplane normal force coefficients in the order of 0.7 were attained at Mach numbers less than 0.9. Buffet frequencies and the variations with Mach number, lift, and angle of attack of buffet intensity are given.
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    Report Date: April 1955


  10. WING-LOAD MEASUREMENTS AT SUPERSONIC SPEEDS OF THE DOUGLAS D-558-II RESEARCH AIRPLANE , Research Memorandum
    Authors: Glenn H. Robinson, George E. Cothren and Chris Pembo
    Report Number: NACA-RM-H54L27
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight measurement of the aerodynamic wing loads on the D-558-II airplane have been made in the Mach number range from 1.0 to 2.0. Results of measurements of the wing-panel normal-force, bending moment, and pitching-moment coefficients, normal-force-curve slope, lateral center of pressure, and chordwise center of pressure are presented.
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    Report Date: March 1955


  11. FLIGHT-DETERMINED PRESSURE DISTRIBUTIONS OVER A SECTION OF THE 35 DEGREE SWEPT WING OF THE DOUGLAS D-558-II RESEARCH AIRPLANE AT MACH NUMBERS UP TO 2.0
    Authors: Gareth H. Jordan and Earl R. Keener
    Report Number: NACA-RM-H55A03
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 29 May 1959
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    Report Date: January 1955
    No. Pages: 38
    Keywords:      Douglas D-558-II aircraft


  12. PRELIMINARY FLIGHT-DETERMINED PRESSURE DISTRIBUTIONS OVER THE WING OF THE DOUGLAS X-3 RESEARCH AIRPLANE AT SUBSONIC AND TRANSONIC MACH NUMBERS
    Authors: Gareth H. Jordan and C. Kenneth Hutchins, Jr.
    Report Number: NACA-RM-H55A10
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Preliminary flight-measured chordwise pressure distributions have been obtained at a wing midsemispan station of the Douglas X-3 research airplane through an angle-of-attack range at Mach numbers of 0.61, 0.78, 0.94, and 1.10. The results of the investigation indicate that the maximum section normal-force coefficient increased from about 0.7 at the lower Mach numbers to about 1.2 at a Mach number of 1.10. The pressure distributions at Mach numbers of about 0.61, 0.78, and 0.94 showed good agreement with wind-tunnel results. At Mach numbers of 0.94 and 1.10 leading-edge flap normal-force and hinge-mount coefficients increased with increase in angle of attack throughout the angle-of-attack range tested and resulted in high normal-force and hinge-moment coefficients at the higher angles of attack.
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    Report declassified: 17 July 1958
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    Report Date: April 1955
    No. Pages: 35
    Keywords:      X-3 aircraft


  13. WING-LOAD MEASUREMENTS OF THE BELL X-5 RESEARCH AIRPLANE AT A SWEEP ANGLE OF 58.70
    Authors: Richard D. Banner, Robert D. Reed and William L. Marcy
    Report Number: NACA-RM-H55A11
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight investigation has been made over an altitude and lift range to determine the wing loads of the Bell X-5 research airplane at a sweep angle of 58.7 degrees at subsonic and transonic Mach numbers. The wing loads were nonlinear over the angle-of-attack range from zero to maximum wing lift. The nonlinear trends were more pronounced at angles of attack above the "pitch-up" where ther is a reduction in the wing lift and an inboard and forward movement in the center of load No apparent effects of altitude on the wing loads were evident from the data obtained in these tests.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 24 February 1956
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    Report Date: April 1955
    No. Pages: 23
    Keywords:      X-5 aircraft


  14. FLIGHT EXPERIENCE WITH TWO HIGH-SPEED AIRPLANES HAVING VIOLENT LATERAL-LONGITUDINAL COUPLING IN AILERON ROLLS
    Authors: Staff
    Report Number: NACA-RM-H55A13
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: During flight tests of two high-speed airplane configurations, violent cross-coupled lateral and longitudinal motions were encountered following abrupt rudder-fixed aileron rolls. The speeds involved ranged from a Mach number of 0.7 to 1.05. The motions were characterized by extreme variations in angles of attack and sideslip which resulted in load factors as large as 6.7g (negative) and 7g (positive) normal acceleration and 2g transverse acceleration.
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    Report declassified: 17 July 1958
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    Report Date: February 1955
    No. Pages: 33


  15. LATERAL STABILITY AND CONTROL CHARACTERISTICS OF THE CONVAIR XF-92 DELTA-WING AIRPLANE AS MEASURED IN FLIGHT
    Authors: Thomas R. Sisk and Duane O. Muhleman
    Report Number: NACA-RM-H55A17
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The lateral stability and control characteristics were investigated on the Convair XF-92A delta-wing airplane during the flights of the NACA research program. The investigation included sideslips, aileron rolls, and rudder pulses at altitudes ranging from 18,000 to 30,000 feet at indicated speeds from 160 to 420 miles per hour. A small amount of data is included with wing fences installed at 60 pecent of the wing semispan for comparison with the basic airplane. The lateral handling characteristics appear satisfactory when viewed in terms of gradually increasing sideslips, lateral control effectiveness, and period, and daming. The pilots objected to the over-all lateral handling characteristics, however, primarily because of the high roll-to-sideslip ratios which probably resulted from the relatively low static directional stability and relatively high effective dihedral. These adverse characteristics were aggravated at low speeds by high adverse yaw and rough air and at high speeds by high airplane response to small controll deflections. The apparent high side force and poor hydraulic control system added to the objectional characteristics. The lateral handling characteristics at low speeds were improved by the installation of wing fences. The improvement apparently resulted from an increase in the static directional stability.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 20 June 1958
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    Report Date: May 1955
    No. Pages: 56
    Keywords:      XF-92A aircraft


  16. FLIGHT DETERMINATION OF THE LONGITUDINAL STABILITY AND CONTROL CHARACTERISTICS OF THE BELL X-5 RESERCH AIRPLANE AT 58.7 DEGREES SWEEPBACK
    Authors: Thomas W. Finch
    Report Number: NACA-RM-H55C07
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The Bell X-5 research airplane has been primarily tested at 58.7 degrees weepback during the program to determine the characteristics of a variable-sweep fighter airplane at transonic speed. Limited stability and control characteristics at 58.7 degrees sweepback have been previously discussed with the presentation of the boundary for reduction of static longitudinal stability at 40,000 feet for Mach numbers up to 0.98. This paper presents the stability and control characteristics in the stable lift range up to Mach numbers near 1.0 at an altitude of 40,000 feet and to slightly lower Mach numbers at altitudes of 25,000 feet and 15,000 feet.
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    Report Date: May 1955
    No. Pages: 31
    Keywords:      X-5 aircraft


  17. FLIGHT MEASUREMENTS OF WING LOADS ON THE CONVAIR XF-92A DELTA-WING AIRPLANE
    Authors: Albert E. Kuhl and Clinton T. Johnson
    Report Number: NACA-RM-H55D12
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight investigation was made at altitudes from 30,000 feet to 35,000 feet to determine the wing loads on the Convair XF-92A airplane over the lift range of the airplane at subsonic and transonic speeds. The theoretical lift-curve slope for a delta wing was calculated and compared with theflight data at a Mach number of 0.75. The wing-panel characteristics display nonlinearities with increasing angle of attack. The wing-panel bending-moment coefficients has nonlinear characteristics throughout the angle-of-attack range, wheras the wing-panel normal-force and pitching-moment coefficients become nonlinear at the higher andgles of attack. In the low-lift region, below the decrease in longitudinal stability, the wing-panel normal-force and pitching-moment coefficients due to angle of attack increase approximately 20 percent of the low-speed values up to a Mach number of 0.83 where the wing reaches its critical Mach number. Above the wing critical Mach number, abrupt changes take place in both parameters with indications of returning to the level of the low-speed values at he highest Mach numbers tested. The lateral center of pressure is located from about 42 percent to 45 percent of the wing-panel semispan for the Mach number range of these tests. The wing-panel normal-force, pitching-moment, and bending-moment coefficients due to elevon deflection, determined in the low-lift region, decrease with increasing Mach number above a Mach number near 0.75.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 6 November 1969
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    Report Date: May 1955
    No. Pages: 35
    Keywords:      XF-92A aircraft


  18. STABILITY AND CONTROL CHARACTERISTICS OBTAINED DURING DEMONSTRATION OF THE DOUGLAS X-3 RESEARCH AIRPLANE
    Authors: Richard E. Day and Jack Fischel
    Report Number: NACA-RM-H55E16
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests were performed with the Douglas X-3 research airplane during the manufacturer's demonstration program and for U. S. Air Force evaluation. These tests covered the Mach number range to 1.21 and an altitude range from 12,800 feet to 34,000 feet. Longitudinal, lateral, and directional stability and control data obtained during these tests in steady flight and maneuvering flight are presented in this paper and are compared with wind-tunnel and rocket-model data. Longitudinal control deflection required to trim the airplane over the Mach number range was generally similar to that of other airplanes, characterized by a stable variation at Mach numbers below 0.92 and a slight nose-down trim change at Mach numbers above 1.07. Data obtained during turns and pull-ups indicated that throughout the Mach number range from 0.65 to 1.21, the apparent static longitudinal stability was positive at low lifts and increased by a factor of about 2 1/2 as Mach number was increased from 0.9 to 1.2. The apparent stability exhibited a gradual decrease as lift increased and mild pitch-ups occurred at Mach numbers above 0.95. The pitch-ups occured at normal-force coefficients of about 0.7 to 0.8, which is slightly below maximum wing lift at a Mach number of approximately 0.95, and about 0.4 to 0.3 below maximum wing lift at Mach numbers greater than 1.0. Difficulty was experienced in performing smooth longitudinal maneuvers. This condition appeared to result from thecombination of control system, pilot, airplane, and their dynamic characteristics; however, additional tests are required to determine the primary cause of the lag and oscillations experienced. Unaccelerated stalls appeaared stable in all configurations tested, except at large angles of attack in the landing configuration where some instability was evident. Roll-off tendencies, which became more severe as the speed was decreased, were apparent in all configurations. Data obtained during sideslips at Mach numbers from 0.84 to 0.98 showed the apparent directional stability to be positive and to increase with increase in Mach number. A smaller degree of apparent stability existed for small angles of sideslip than existed for largerangles. Meager aileron effectiveness data obtained at Mach numbers of 0.89 to 0.98 indicated that the control effectiveness was generally linear with deflection and exhibited little change with increase in Mach number. Comparison of flight data with wind-tunnel and rocket-model tests showed similar trends and good quantitative agreement.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 17 July 1958
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    Report Date: July 1955
    No. Pages: 52
    Keywords:      X-3 aircraft


  19. FLIGHT EXPERIENCE OF INERTIA COUPLING IN ROLLING MANEUVERS
    Authors: Joseph Weil, Ordway B. Gates, Jr., Richard D. Banner and Albert E. Kuhl
    Report Number: NACA-RM-H55E17B
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Violent coupled lateral-longitudinal motions have been encountered in flight on two airplanes during abrupt aileron rolls at relatively high speed. During these motions, variouss structural design loads and load factors were either exceeded or approached. It was demonstrated on one airplane that the motions can be approximated reasonably well by using a five-degree-of-freedom analysis. From flight tests of the swept-wing airplane at relatively high altitude, it was found that the severity of the divergent tendency increased with roll velocity and was sensitive to roll direction and stabilizer input. Calculated results indicated that considerably more critical conditions from the loads standpoint can be expected at lower altitudes when the roll is initiated from a pull-up condition. Perhaps one of the fundamental reasons for the occurrence of the large motions on both airplanes was the presence of insufficient directional stability. Doubling the directional stability level of the swept-wing airplane resulted in substantioally improved flight characteristics; but calculations indicated that, if the tail size is increased beyond a certain point, considerably higher tail loads and larger peak normal accelerations can be obtained than with a tail affording a somewhat lower level of stability. At present, analytical investigations are under way to enable a better understanding of the overall problem of coupled lateral-longitudinal motions in rolling maneuvers. It is not yet known whether a proctical design approach exists that would produce desirable characteristics for a large range of flight conditions without the sacrifice of performance or the resort to artificial stabilization. It is also true that coupling lateral and longitudinal degrees of freedom should be considered for load evaluations of rolling maneuvers on most high-speed airplanes. There is a deterioration in the static directional stability of many contemporary designs at the higher angles of attack and sideslip, and also with increase in supersonic Mach number, that can and have procuced violent motions in flight. Recently at the NACA High-Speed Flight Station, some rather violent coupled lateral-longitudinal motions have been experienced during abrupt aileron rolls on several airplanes in which a level of directional stability was present that would probably have been deemed acceptable for previous airplanes. Because this flight experience should be of considerable general interest to the loads engineer, inasmuch as it obviously affects the determination of design loads, it is believed timely to review briefly the problem and indicate some of the factors affecting its severity.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 17 July 1958
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    Report Date: July 1955
    No. Pages: 13


  20. FLIGHT MEASUREMENTS OF HORIZONTAL-TAIL LOADS ON THE BELL X-5 RESEARCH AIRPLANE AT A SWEEP ANGLE OF 58.7 DEGREES
    Authors: Robert D. Reed
    Report Number: NACA-RM-H55E20A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight investigation was made at altitudes of 40,000, 25,000 and 15,000 feet to determine the horizontal-tail loads of the Bell X-5 research airplane at a sweep angle of 58.7 degrees over the lift range of the airplane for Mach numbers from 0.61 to 1.00. The horizontal-tail loads were found to be nonlinear with lift throughout the lift ranges tested at all Mach numbers except at a Mach number of 1.00. The balancing tail loads reflected the changes which occur in the wing characteristics with increasing angle of attack. The nonlinearities were, in general, more pronounced at the higher angles of attack near the pitch-up where the balancing tail loads indicate that the wing-fuselage combination becomes unstable. No apparent effects of altitude on the balancing tail loads were evident over the comparable lift ranges of these tests at altitudes from 40,000 feet to 15,000 feet. Comparisons of balancing tail loads obtained from flight and wind-tunnel tests indicated discrepancies in absolute magnitudes, but the general trends of the data agree. Some differences in absolute magnitude may be accounted for by the tail load carried inboard of the strain-gage station and the load induced on the fuselage by the presence of the tail. These loads were not measured in flight.
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    Report Date: July 1955
    No. Pages: 20
    Keywords:      X-5 aircraft


  21. FLIGHT MEASUREMENTS OF THE LATERAL RESPONSE CHARACTERISTICS OF THE CONVAIR XF-92A DELTA-WING AIRPLANE
    Authors: Euclid C. Holleman
    Report Number: NACA-RM-H55E26
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: As part of the flight reserch program conducted with the Convair XF-92A delta-wing research arplane, rudder pulse maneuvers were obtained at an altitude of about 30,000 feet over a Mach number range of 0.52 to 0.92. Tests were made with and without a wing fence. By analyzing these maneuvers the characteristics of the airplane transient, airplane stability derivatives, and frequency-responce characteristics were measured. The airplane handling qualities were improved by the addition of wing fences. The agreement between experimental and calculated stability derivatives was fair to poor. However by using transfer-function equations from the lateral equations of motion and the experimental stability derivatives, frequency responses were calculated that compared favorably with those determined by Fourier transformation.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 16 May 1958
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    Report Date: August 1955
    No. Pages: 38
    Keywords:      XF-92A aircraft


  22. EFFECT OF SEVERAL WING MODIFICATIONS ON THE LOW-SPEED STALLING CHARACTERISTICS OF THE DOUGLAS D-558-II RESEARCH AIRPLANE
    Authors: Jack Fischel and Donald Reisert
    Report Number: NACA-RM-H55E31A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The low-speed stalling and lift characteristics of the Douglas D-558-II airplane were measured in a series of 1 g stall approaches performed with several wing modifications designed to alleviate swept-wing instability and pitch-up. The airplane configurations investigated include the basic wing configuration and two wing-fence configurations in conbination with retracted, free-floating, or extended slats, and a wing leading-edge chord-extension configuration. All configurations were investigated with flaps and landing gear retracted and extended at an altitude of about 20,000 feet. With stats, flaps, and landing gear retracted, none of the wing modifications investigated had an appreciable effect on the lift or stability characteristics at low and moderate angles of attack. Regardless of wing-fence configuration, appreciably larger values of peak normal-force coefficient were attained with stats unlocked (free floating) or fully extended than with slats closed. Wing fences and the chord-extension tended to delay the onset of instability with slats retracted, and the stable region was further extended for the configurations with either no wing fences or inboard wing fences when the slats were free floating or extended. With flaps and landing gear extended, only the fully extended slat configuration affected the variation of normal-force coefficient with angle of attack by increasing this variation slightly. Peak values of normal-force coefficient attained were the same for all configurations except the chord-extension configuration. For this configuration excessive buffeting caused earlier termination of the maneuver. Most of the configurations had little or no effect on the stability characteristics over most of the lower and moderate angle-of-attack range. The airplane appeared somewhat more stable, however, with no wing fences installed than with wing fences intalled when the slats were extended. At larger angles of attack and with slats extended, inboard wingfences materially improved the stability caharacteristics of the airplane. At any given angle of attack, wing flaps provided an increment in normal-force coefficient of about 0.3; whereas, the free-floating or fully extended slats provided zero incremental lift except at very large angles of attack. The airplane generally appeared more stable longitudinally at comparable speeds with flaps deflected than with flaps retracted, but marginal dynamic lateral stability was evident for several configurations with flaps extended or retrated. In general, acequate stall warning in the form of buffeting was noted by the pilot in the stable region of flight, particularly for the chord-extension configuration for which buffeting appeared aggravated.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 24 June 1958
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    Report Date: July 1955
    No. Pages: 63
    Keywords:      Douglas D-558-II aircraft


  23. WING PRESSURE DISTRIBUTIONS OVER THE LIFT RANGE OF THE CONVAIR XF-92A DELTA-WING AIRPLANE AT SUBSONIC AND TRANSONIC SPEEDS
    Authors: Earl R. Keener and Gareth H. Jordan
    Report Number: NACA-RM-H55G07
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Chordwise pressure distributions were measured over the left wing of the Convair XF-92A delta-wing airplane to determine the effect of lift upon the wing characteristics at subsonic and transonic Mach numbers. The data were obtained throughout the Mach number range of 0.30 to 0.93. Reynolds number based on the mean aerodynamic chord of the wing varied between 22x10(to the 6th power) and 49x10 (to the 6th power). High leading-edge suction, followed by a leading-edge-separation vortex, resulted from the small leading-edge radius. The outboard wing sections experienced a higher effective angle of attack because of the high degree of taper and large degree of sweep. As a result the high leading-edge suction and the leading-edge-separation vortex occurred first at the wing tip and moved inboard with increasing angle of attack. The critical Mach number of the wing decreased from 0.82 to 0.65 with an increase in airplane normal-force coefficient from 0.06 to 0.17. The wing-section lift and pitching-moment curves were very nonlinear with angle of attack because of the large three-dimensional effects. The wing sections at the tip stalled at an airplane normal-force coefficient of about 0.20. The spanwise load distributions tended to approach a tringular shape at high lift as a result of the inboard movement of wing-section stall. Wing-section stall resulted in a local reversal in the elevon-section load from down-load to up-load.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 17 July 1958
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    Report Date: November 1955
    No. Pages: 135
    Keywords:      XF-92A aircraft
    Notes: Unclassified July 17, 1958


  24. BEHAVIOR OF THE BELL X-1A RESEARCH AIRPLANE DURING EXPLORATORY FLIGHTS AT MACH NUMBERS NEAR 2.0 AND AT EXTREME ALTITUDES
    Authors: Hubert M. Drake and Wendell H. Stillwell
    Report Number: NACA-RM-H55G25
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight program has been conducted by the U. S. Air Force consisting of exploratory flights to determine the Mach number and altitude capabilities of the Bell X-1A research airplane. On two flights of the X-1A airplane, one reaching a Mach number of about 2.44, the other a geometric altitude of about 90,000 feet, lateral stability difficulties were encountered which resulted in uncontrolled rolling motions of the airplane at Mach numbers near 2.0. Analysis indicates that this behavior apparently results from a combination of low directional stability and damping in roll and may be aggravated by high control friction and rocket motor misalignment. The deterioration of directional stability with increasing Mach number can lead to severe longitudinal-lateral coupling at low roll rates. The misalignment of the rocket motor could induce sufficiently high roll velocities to excite these coupled motions. Adequate control of these motions was virtually impossible because of the high control friction. In the absence of rolling, poor lateral behavior might be expected at somewhat higher Mach numbers because wind-tunnel data indicate neutral directional stability at about M = 2.35.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 25 January 1960
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    Report Date: September 1955
    No. Pages: 25
    Keywords:      X-1A aircraft


  25. FLIGHT MEASUREMENTS OF DIRECTIONAL STABILITY TO A MACH NUMBER OF 1.48 FOR AN AIRPLANE TESTED WITH THREE DIFFERENT VERTICAL TAIL CONFIGURATIONS
    Authors: Hubert M. Drake, Thomas W. Finch and James R. Peele
    Report Number: NACA-RM-H55G26
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests have been performed to measure the directional stability of a fighter-type airplane over the Mach number range from 0.72 to 1.48. The tests were made at altitudes of 40,000 feet and 30,000 feet and employed three different vertical tails of varying aspect ratio or area, or both. These tests showed that the directional stability for all tail configurations increased with an increase in tail aspect ratio or area, or both, over the entire Mach number range and decreased with increasing supersonic Mach number above 1.15.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 17 July 1958
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    Report Date: January 1955
    No. Pages: 24


  26. FLIGHT EXPERIENCE WITH A DELTA-WING AIRPLANE HAVING VIOLENT LATERAL-LONGITUDINAL COUPLING IN AILERON ROLLS
    Authors: Thomas R. Sisk and William H. Andrews
    Report Number: NACA-RM-H55H03
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: During a flight investigation of the lateral stability characteristics of a high-speed delta-wing airplane, violent cross-coupled lateral and longitudinal motions were encountered. The maneuver which produced these motions was an abrupt, rudder-fixed aileron roll performed at a Mach number of 0.75 at about 40,000 feet. The motions were characterized by extreme variations in angle of attack and angle of sideslip which caused the airplane to exceed the normal and transverse acceleration limitations.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 8 December 1961
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    Report Date: October 1955
    No. Pages: 19


  27. FLIGHT MEASUREMENTS OF THE DYNAMIC LATERAL AND LONGITUDINAL STABILITY OF THE BELL X-5 RESEARCH AIRPLANE AT 58.7 DEGREES SWEEPBACK
    Authors: Edward N. Videan
    Report Number: NACA-RM-H55H10
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An investigation has been made of the dynamic stability of the Bell X-5 research airplane at 58.7 degrees sweepback and at altitudes of 40,000 feet and 25,000 feet over a Mach number range of 0.50 to 0.97. The results of this investigation show that the longitudinal oscillatory motions are well damped over the entire Mach number range except for residual oscillations resulting principally from engine gyroscopic coupling with the lateral oscillatory mode. The lateral oscillatory mode is moderately will damped except for nonlinear damping characteristics above a Mach number of 0.80. The damping is high for large amplitudes but for sideslip angles of less than 2 degrees the damping is poor. The airplane is highly sensitive to small inadvertent control motions which produce apparently undamped small amplitude oscillations in the Mach number range above M = 0.80. By U.S. Military Specifications, the lateral oscillation is marginal except in the nonlinear damping range where it is definitely unsatisfactory. The longitudinal and lateral frequency-response characteristics were obtained and some coupling effects were noted. Longitudinal frequency-response calculations were characterized by adjacently located double peaks in amplitude ratio. This behavior is satisfactorily explained by theoretical calculations of frequency response involving engine gyroscopic coupling to the lateral mode. Lateral frequency responses showed some dependence on the direction of the initial disturbance.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 24 February 1958
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    Report Date: October 1955
    No. Pages: 70
    Keywords:      X-5 aircraft


  28. FLIGHT MEASUREMENTS OF THE VERTICAL-TAIL LOADS ON THE CONVAIR XF-92A DELTA-WING AIRPLANE
    Authors: Clinton T. Johnson
    Report Number: NACA-RM-H55H25
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The aerodynamic loads acting over the vertical tail were determined from steady and maneuvering flight during the investigation of the lateral stability and control characteristics of the Convair XF-92A airplane. The results presented in this paper were obtained from rudder pulses and gradually increasing sideslips over the Mach number range from 0.50 to 0.87 at altitudes between 30,000 feet and 20.000 feet. The vertical-tail panel bending-moment and normal-force characteristics are essentially linear with indreasing sideslip angle both in rudder-fixed and trimmed maneuvers. A comparison of the bending-moment and normal-force parameters derived from rudder-fixed oscillations and the corresponding parameters derived from gradual maneuvers indicates similar trends with Mach number. The effect of rudder deflection is to reduce the slope of the vertical-tail normal-force-coefficient variation with sideslip angle and to move the lateral location of the center of pressure of the additional air load inboard about 5 percent of the span of the vertical-tail panel. The vertical-tail bending-moment and normal-force coefficients resulting from rudder deflections are essentially constant below a Mach number of 0.80 with an apparent tendency for both parameters to increase at the higher Mach numbers tested.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 11 February 1957
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    Report Date: October 1955
    No. Pages: 24


  29. MACH NUMBER MEASUREMENTS AND CALIBRATIONS DURING FLIGHT AT HIGH SPEEDS AND AT HIGH ALTITUDES INCLUDING DATA FOR THE D-558-II RESEARCH AIRPLANE
    Authors: Cyril D. Brunn and Wendell H. Stillwell
    Report Number: NACA-RM-H55J18
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The accuracies associated with Mach number measurements based on pitot-static pressures are evaluated with respect to the individual errors of the measuring instruments, recording instruments, position-error calibrations, and time lag. The methods and instruments employed at the National Advisory Committee for Aeronautics High-Speed Flight Station at Edwards, Calif. are used to illustrate the magnitudes of the errors associated with Mach number measurements at supersonic speeds and at high altitudes. It is indicated that the largest Mach number errors are caused by inaccuracies of positon-error calibrations and that an over-all accuracy of 1 pecent will be difficult to attain. Data are presented for two modified methods of the basic radar phototheodolite method of position-error calibration. These methods are based on radiosonde measurements of pressure and temperature and are not dependent on a fly-by or low-level position-error calibration. The practical application of the instruments and methods employed by the NACA is given for the maximum altitude and maximum Mach number flights of the D-558-II research airplane in which a maximum altitude of 83,235 feet, a maximum Mach number of 2.005, and a maximum true airspeed of 1,291 miles per hour were attained.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 24 June 1958
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    Report Date: March 1955
    No. Pages: 32
    Keywords:      Douglas D-558-II aircraft


  30. FLIGHT INVESTIGATION OF THE EFFECT OF VERTICAL-TAIL SIZE ON THE ROLLING BEHAVIOR OF A SWEPT-WING AIRPLANE HAVING LATERAL-LONGITUDINAL COUPLING
    Authors: Thomas W. Finch, James R. Peele and Richard E. Day
    Report Number: NACA-RM-H55L28A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests have been performed over a Mach number range of 0.73 to 1.39 to determine the rolling behavior of a swept-wing airplane having lateral-longitudinal coupling. The tests were made at altitudes of 40,000 feet and 30,000 feet and employed three different vertical tails with varying aspect ratio or area, or both, and two wing configurations, the basic wing, and the basic wing plus wing-tip extensions. The airplane with the original vertical tail exhibited violent motions resulting from cross coupling at the higher rolling velocities. Consequently, tests with this configuration were limited to low aileron deflections and to bank angles less than 90 detgees. Doubling the directional stability by increasing the tail area 27 percent and the tail aspect ratio 32 percent greatly improved the rolling behavior enabling rolling rates on the order of 3 to 4 radians per second to be obtained. The adverse sidesilp encountered during roll maneuvers decreased with increasing speeds to neglibible values over a Mach number range of approximately 1.00 to 1.05; the sideslip then increased in a favorable direction with further increases in speed. The present allowable sideslip angles imposed by structural limitations were not approached at either subsonic or supersonic speeds. Engine gyroscopic effects caused the rolling behavior to be worse in left rolls at subsonic speeds (adverse yaw) and in right rolls at supersonic speeds (favorable yaw). Small airplane nose-up stabilizer motion during the roll made the behavior considerably worse at subsonic speeds, whereas small stabilizer motion in the opposite direction improved the behavior. At supersonic speeds the reverse is true, but to a lesser degree.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 10 February 1959
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    Report Date: April 1955
    No. Pages: 34


  31. SOME RECENT RESEARCH ON THE HANDLING QUALITIES OF AIRPLANES
    Authors: Walter C. Williams and William H. Phillips
    Report Number: NACA-RM-H55L29A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Results of recent research on the handling qualities of airplanes are reiewed. Among the subjects considered are dynamic longitudinal stability, transonic trim changes, pitch-up due to decreasing airspeed, dynamic lateral stability, aileron control, rudder control, and mechanical characteristics of power control systems.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 8 November 1957
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    Report Date: February 1955
    No. Pages: 20
    Keywords:      Handling qualities


  32. FLIGHT CALIBRATION OF FOUR AIRSPEED SYSTEMS ON A SWEPT-WING AIRPLANE AT MACH NUMBERS UP TO 1.04 BY THE NACA RADAR-PHOTOTHEODOLITE METHOD
    Authors: Jim Rogers Thompson, Richard S. Bray and George E. Cooper
    Report Number: NACA-TN-3526
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The calibrations of four airspeed systems installed in a North American F-86A airplane have been determined in flight at Mach numbers up to 1.04 by the NACA radar-phototheodolite method. The variation of the static-pressure error per unit indicated impact pressure is presented for three systems typical of those currently in use in flight research, a nose boom and two different wing-tip booms, and for the standard service system installed in the airplane. A limited amount of information on the effect of airplane normal-force coefficient on the static-pressure error is included. The results are compared with available theory and with results from wind-tunnel tests of the airspeed heads alone. Of the systems investigated, a nose-boom installation was found to be most suitable for research use at transonic and low supersonic speeds because it provided the greatest sensitivity of the indicated Mach number to a unit change in true Mach number at very high subsonic speeds, and because it was least sensitive to changes in airplane normal-force coefficient. The static-pressure error of the nose-boom system was small and constant above a Mach number of 1.03 after passage of the fuselage bow shock wave over the airspeed head.
    Distribution/Availability: Unclassified - Unlimited
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    Report Date: November 1955
    No. Pages: 37
    Keywords:      F-86A aircraft
 
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