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  1. STALL CHARACTERISTICS OBTAINED FROM FLIGHT 10 OF NORTHROP X-4 NO. 2 AIRPLANE (USAF NO. 46-677)
    Authors: Melvin Sadoff and Thomas R. Sisk
    Report Number: NACA-RM-A50A04
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NACA instrumentation has been installed in the X-4 airplanes to obtain stability and control data during the acceptance tests conducted by the Northrop Aircraft Corporation. This report presents data obtained on the stalling caracteristics of the airplane in the clean and gear-down configurations. The center of gravity was located at approximately 18 percent of the mean aerodynamic chord during the tests. The results indicated that the airplane was not completely stalled when stall was gradually approached during nominally unaccelerated flight but that it was completely stalled during a more abruptly approached stall in accelerated flight. The stall in accelerated flight was relatively mild, and this was attributed to the nature of the variation of lift with angle of attack for the 0010-64 airfoil section, the plan form of the wing, and to the fact that the initial sideslip at the stall produced (as shown by wind-tunnel tests of a model of the airplane) a more symmetrical stall pattern.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 16 July 1951
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    Report Date: February 1950
    No. Pages: 23


  2. LONGITUDINAL-STABILITY CHARACTERISTICS OF THE NORTHROP X-4 AIRPLANE (USAF NO. 46-677)
    Authors: Melvin Sadoff and Thomas R. Sisk
    Report Number: NACA-RM-A50D27
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The results obtained from several recent flights on the Northrop X-4 No. 2 airplane are presented. Information is included on the longitudinal-stability cahracteristics in straight flight over a Mach number range of 0.38 to about 0.63, the longitudinal-stability characteristics in accelerated flight over a Mach number range of 0.43 to about 0.79, and the short-period longitudinal-oscillation characteristics at Mach numbers of 0.49 and 0.78. It was shown that the stick-fixed and stick-free static longitudinal stability, as measured in straight flight, were positive over the test speed range with the center of gravity located at about 18.0 percent of the mean aerodynamic chord. During the longitudinal-stability tests in accelerated flight an inadvertent pitch-up of the airplane occurred at a Mach number of about 0.79 and a normal-force coefficient of about 0.45 (normal acceleration factor, the ratio of the net aerodynamic force along the airplane Z axis to the weight of the airplane = 5), in which the acceleration built up rapidly to the ratio of the net aerodynamic force along the airplane Z axis to the weight of the airplane = 6.2 (which was in excess of the load factor, 5.2, required for demonstration of the airplane) before recovery could be initiated. A comparison of the experimentally determined elevon angles required for balance and the elevon-angle gradients with values estimated from limited wind-tunnel data showed fairly good agreement. Wind-tunnel data, however, were not available in the region where the pitch-up occurred so that an evaluation in this regard was not possible. The short-period oscillation was lightly damped and did not meet the Air Force requirements for satisfactory handling qualities. The pilot, however, did not object to the low damping characteristics of this airplane for small-amplitude oscillations. Theory predicted the period of the short-period longitudinal oscillation fairly well; however, the damping evaluated from the theory indicated considerably greater damping than was actually measured in flight, especially at the higher Mach numbers.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 1 December 1955
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    Report Date: June 1950
    No. Pages: 23


  3. SUMMARY REPORT OF RESULTS OBTAINED DURING DEMONSTRATION TESTS OF THE NORTHROP X-4 AIRPLANES
    Authors: Melvin Sadoff and Thomas R. Sisk
    Report Number: NACA-RM-A50I01
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Results obtained during the demonstration flight tests of the Northrop X-4 No. 1 and No. 2 airplanes are presented. Information is included on the static and dynamic longitudinal- and lateral-stability characteristics, the stalling characteristics, and the buffet boundary. The data indicated that the airplane was almost neutrally stable in straight flight at low Mach numbers with the center of gravity located at about 21.4 percent of the mean aerodynamic chord for the clean configuration. In accelerated flight over a Mach number range of about 0.44 to 0.84 the airplane was longitudinally stable up to a normal-force coefficient of about 0.4. At higher values of normal-force coefficient and at the higher (aproximately Mach 0.8) Mach numbers a longitudinal instability was experienced. The X-4 airplane does not satisfy the Air Force specifications for damping of the short-period longitudinal oscillation. The pilot, however, did not consider the low ddamping characteristics of the airplane objectionable for small disturbances. An objectionable undamped oscillation about all three axes was experienced, however, at the highest test Mach number of 0.88. Theory predicted the period of the short-period longitudinal oscillation fairly well, while, in general, the theoretical damping indicated a higher degree of stability than was actually experienced. This discrepancy was traced to a considerable error in the estimation of the rotational damping factor. The directional stability of the X-4 airplane as measured in steady sideslips was high and essentially constant over the speed range covered, while the dihedral effect decreased considerably with an increase in airspeed. The damping of the lateral oscillation does not meet the Air Force requirements for satisfactory handling qualities over the Mach number range covered. The data indicated decreased damping as the flight Mach number was increased above about 0.5, and at high Mach numbers (M>0.8) and at high altitudes the X-4, in common with other transonic research airplanes, experienced a small amplitude undamped lateral oscillation. The dynamic lateral-stability characteristics were estimated fairly well by theory at low Mach numbers and at a pressure altitude of 10,000 feet. At 30,000 feet, however, at Mach numbers above about 0.6, the theory again indicated a higher degree of stability than was actually obtained. For the conditions covered in these tests the stalling characteristics of the X-4 airplane, as measured in stall approaches in straight flight and in an accelerated stall to about 1.6g, were, in general, satisfactory. Both the stall approaches and the stall were characterized by a roll-off to the right. The X-4 buffet boundary showed a sharp drop-off in the normal-force coefficient for the onset of buffeting as the flight Mach number exceeded 0.8. The boundary was almost identical to that obtained for the D-558-II research airplane at comparable Mach numbers.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 8 September 1954
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    Report Date: December 1950
    No. Pages: 45


  4. FLIGHT MEASUREMENTS WITH THE DOUGLAS D-558-II (BUAERO NO. 37974) RESEARCH AIRPLANE STATIC LATERAL AND DIRECTRIONAL STABILITY CHARACTERISTICS AS MEASURED IN SIDESLIPS AT MACH NUMBERS UP TO 0.87
    Authors: S. A. Sjoberg
    Report Number: NACA-RM-L50C14
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight measurements were made in sideslips of the static lateral and directional stability characteristics of the Douglas D-558-II (BuAero No. 37974) research airplane. The directional stability of the airplane was positive in both the clean and landing conditions at all test speeds. About 2 degrees of rudder deflection were required to produce 1 degree of sideslip in both the clean and landing conditions. There was no decrease in the effectiveness of the rudder in producing sideslip up to the highest Mach number reached (0.87).
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 29 May 1957
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    Report Date: May 1950
    No. Pages: 28


  5. FLIGHT MEASUREMENTS WITH THE DOUGLAS D-558-II (BUAERO NO. 37974) RESEARCH AIRPLANE. DETERMINATION OF THE AERODYNAMIC CENTER AND ZERO-LIFT PITCHING-MOMENT COEFFICIENT OF THE WING-FUSELAGE COMBINATION BY MEANS OF TAIL-LOAD MEASUREMENTS IN THE MACH NUMBER RA , Research Memorandum
    Authors: John P. Mayer, George M. Valentine and Geraldine C. Mayer
    Report Number: NACA-RM-L50D10
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Determination of the aerodynamic center and zero-lift pitching-moment coefficient of the wing-fuselage combination by means of tail-load measurements in the Mach number range from 0.37 to 0.87.
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    Report Date: July 1950


  6. FLIGHT INVESTIGATION OF THE AILERON CHARACTERISTICS OF THE DOUGLAS D-558-I AIRPLANE (BUAERO NO. 37972) AT MACH NUMBERS BETWEEN 0.6 AND 0.89
    Authors: Jim Rogers Thompson, William S. Roden and John M. Eggleston
    Report Number: NACA-RM-L50D20
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Abrubt, rudder-fixed aileron rolls have been made with the Douglas D-558-I airplane (BuAero No. 37972) at Mach numbers between 0.6 and 0.89. Rolls were made at aileron deflections between one-eighth and one-half the maximum available deflection. The results obtained indicate that the aileron effectiveness is independent of Mach number and defliection within the range investigated. Limited information on the lateral trim and handling qualities of the airplane at high Mach numbers is presented.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 9 September 1959
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    Report Date: May 1950
    No. Pages: 21


  7. FLIGHT MEASUREMENTS WITH THE DOUGLAS D-558-II (BUAERO NO. 37974) RESEARCH AIRPLANE LATERAL CONTROL CHARACTERISTICS AS MEASURED IN ABRUPT AILERON ROLLS AT MACH NUMBERS UP TO 0.86
    Authors: J.V. Wilmerding, W.H. Stillwell and S.A. Sjoberg
    Report Number: NACA-RM-L50E17
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight measurements were made of the lateral control characteristics of the Douglas D-558-II airplane in abrupt rudder-fixed aileron rolls. In the Mach number range from 0.50 to 0.86 the aileron rolling effectiveness is substantially constant and the rate of change of the maximum wing-tip helix angle with total aileron deflection (rate of change of maximum wing-tip helix angle with total aileron deflection, radians per degree)has a value of 0.0027 radian per degree. Extraploated data indicate that in this Mach number range full aileron deflection of 30 degrees will produce a maximum wing-tip helix angle pb/2V of about 0.08 radian. As the speed is reduced below a mach number of 0.50 a marked decrease occurs in the maximum value of pb\2V obtainable with a given aileron deflection. This decrease occurs because the dihedral effect increases with decrease in speed and the adverse sideslip angles reached in the rolls at low speed are larger. At an indicated airspeed of 150 miles per hour in the landing condition, full aileron deflection will produce a maximum pb/2V of 0.04 radian, which for standard sea-level conditions corresponds to a rolling velocity of 40 degrees per second. In the opinion of the pilots this rolling velocity is sufficiently high for the landing condition with this airplane. It is the opinion of several NACA pilots that the maximum usable rolling velocity is on the order of 2.5 radians per second. In the Mach number range from 0.42 to 0.86 at an altitude of 15,000 feet rolling velocites greater than 2.5 radians per second can be obtained with less than full aileron deflection. The data indicate that in going from high to low lift coefficient the yawing moment due to rolling changes direction. At high lift coefficients the sideslip due to roll is in the same direction as the roll (right roll produces right sideslip), but at low lift coefficients the opposite tendency is present.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 29 May 1957
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    Report Date: July 1950
    No. Pages: 25


  8. FLIGHT MEASUREMENTS WITH THE DOUGLAS D-558-II BUAERO NO. 37974 RESEARCH AIRPLANE. MEASUREMENTS OF THE BUFFET BOUNDARY AND PEAK AIRPLANE NORMAL FORCE COEFFICIENTS AT MACH NUMBERS UP TO 0.90 , Research Memorandum
    Authors: John P. Mayer and George M. Valentine
    Report Number: NACA-RM-L50E31
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Measurements of the buffet boundary and peak airplane normal force coefficients at Mach numbers up to 0.90.
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    Report Date: August 1950


  9. EFFECT OF AN AUTOPILOT SENSITIVE TO YAWING VELOCITY ON THE LATERAL STABILITY OF THE DOUGLAS D-558-II AIRPLANE , Research Memorandum
    Authors: Ordway B. Gates and Leonard Sternfield
    Report Number: NACA-RM-L50F22
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A theoretical investigation has been made to determine the effect on the lateral stability of the Douglas D-558-II airplane of an autopilot sensitive to yawing velocity. The effects of inclination of the gyro spin axis to the flight path and of time lag in the autopilot were also determined. The flight conditions investigated included landing at sea level, approach condition at 12,000 feet, and cruising at 12,000 feet at Mach numbers of 0.80 and 1.2. The results of the investigation indicated that the lateral stability characteristics of the D-558-II airplane for the flight condition discussed should satisfy the Air Force - Navy period-damping criterion when the proposed autopilot is installed. Airplane motions in sideslip subsequent to a disturbance in sideslip are presented for several representative flight conditions in which a time lag in the autopilot of 0.10 second was assumed.
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    Report Date: August 1950


  10. FLIGHT MEASUREMENTS WITH THE DOUGLAS D-558-II (BUAERO NO. 37974) RESEARCH AIRPLANE LOW-SPEED STALLING AND LIFT CHARACTERISTICS
    Authors: W. H. Stillwell, J. V. Wilmerding and R. A. Champine
    Report Number: NACA-RM-L50G10
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The low-speed stalling and lift characteristics of the D-558-II airplane were measured in a series of 1 g stalls in four different airplane configurations. With the slats locked closed and the flaps up or down, the airplane was unstable at angles of attack greater than about 9 degrees. With the flaps up this corresponds to a normal-force coefficient of about 0.8 and with the flaps down, about 1.07. Because of this instability, the airplane tended to pitch to high angles of attack; at these high angles of attack, violent rolling and yawing motions sometimes occurred. In one case with the flaps down and the slats locked the airplane went into a spin after pitching up to high angles of attack. The pilots considered the stalling characteristics of the airplane with the slats locked to be very objectionable. No data are presented in this paper on the stalling characteristics in maneuvering flight, but the pilots considered the longitudinal instability particularly objectionable in maneuvering flight. With the slats unlocked and the flaps up or down the airplane was unstable at angles of attack greater than about 23 degrees. Uncontrolled-for rolling and yawing motions due to stalling were present when the airplane was unstable in the high angle-of-attack range. With the slats unlocked and the flaps and landing gear up or down, there was adequate stall warning in the form of buffeting and lateral oscillations of the airplane. With the slats locked, slight buffeting of the airplane occurred at a normal-force coefficient slightly less than the normal-force coefficient at which the airplane became longitudinally unstable. With the flaps up and the slats locked, the highest normal-force coefficient obtained was 1.13 at an angle of attack of about 17.5 degrees. The highest normal-force coefficient obtained with the flaps up and the slats unlocked was 1.46 at and angle of attack of 36 degrees, and in the angle-of-attack range from 23 degrees to 30 degrees the normal-force coefficient had a substantially constant value of 1.32. At the lower angles of attack with the slats locked or unlocked deflecting the flaps produced an increment in normal-force coefficient at a given angle of attack of about 0.26. The highest normal-force coefficient obtained with the flaps down and the slats locked or unlocked was about 1.65. This value was attained at an angle of attack of about 35.5 degrees with the slats locked and at an angle of attack of about 38 degrees with the slats unlocked. However, in the angle-of-attack range from 12 degrees to 32 degrees considerably greater normal-force coefficients were obtained with the slats unlocked than with the slats locked.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 29 September 1950
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    Report Date: September 1950
    No. Pages: 38


  11. ELEVATOR-STABILIZER EFFECTIVENESS AND TRIM OF THE X-1 AIRPLANE TO A MACH NUMBER OF 1.06
    Authors: Hubert M. Drake and John R. Carden
    Report Number: NACA-RM-L50G20
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Limited measurements of elevator-stabilizer effectiveness and trim of the X-1 airplane with the 10-percent-thick wing and 8-percent-thick tail have been presented previously to a Mach number of about 0.93. Subsequent flights have permitted refinement and extension of these data to higher Mach numbers. The data presented in this report were obtained at about 40,000 feet altitude at Mach numbers between 0.78 and 1.06 for normal-force coefficients between 0.26 and 0.42. The data show that at Mach numbers between 0.78 and 0.92, the variation of elevator position is gradual for all the stabilizer settings tested. Above a Mach number of about 0.92, trim changes are more pronounced. The magnitude and direction of these trim changes and the Mach number at which they occur change with stabilizer incidence. The data indicate that stabilizer angles of 2 degrees and 0.5 degrees are the limit settings for which the airplane can be trimmed with the elevator alone through the Mach number range up to M = 1.0. Because of the high altitude of flight the stick forces involved were moderate, maximum values of 30 pounds pull and 50 pounds push being obtained. The relative elevator-stabilizer effective rate of change of stabilizer incidence angle, with respect to elevator angle, decreases from a value of 0.25 at a Mach number of 0.78 to a minimum value of 0.05 at Mach number of 1.0. At Mach numbers between 1.01 and 1.06 the effectiveness increases. The variation of elevator deflection with stabilizer incidence was nonlinear between Mach numbers of 0.94 and 0.97. The variation of rate of change of stabilizer incidence angle, with respect to elevator angle, with Mach number and the nonlinearity of this curve at Mach numbers between 0.94 and 0.97 were primarily responsible for the difference between the trim curves obtained at the various stabilizer settings. It was found that, with the elevator fixed at zero, only about 0.5 degrees of stabilizer movement would be required to trim through the Mach number range from 0.78 to 1.02 but greater movements would be required at Mach numbers above 1.02.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 19 July 1956
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    Report Date: November 1950
    No. Pages: 13


  12. FLIGHT MEASUREMENTS OF THE PRESSURE DISTRIBUTION ON THE WING OF THE X-1 AIRPLANE (10-PERCENT-THICK WING) OVER A CHORDWISE STATION NEAR THE MIDSPAN, IN LEVEL FLIGHT AT MACH NUMBERS FROM 0.79 TO 1.00 AND IN A PULL-UP AT A MACH NUMBER OF 0.96
    Authors: H. Arthur Carner and Ronald J. Knapp
    Report Number: NACA-RM-L50H04
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Measurements of the chordwise pressure distribution over the 10-percent-thick wing of the X-1 research airplane have been made at a section near the midspan of the left wing. Data presented are for a Mach number range from 0.79 to 1.00 at a section normal-force coefficient of about 0.32 and for section normal-force coefficients up to 1.00 at a Mach number of approximately 0.96. The results show that the section center of load moves aft from about 32 percent chord at Mach number 0.79 to 40 percent chord at Mach number 0.84, and then forward to 18 percent chord at Mach number 0.89. The section center of load moves aft to 45 percent chord at Mach number 0.95 and then remains approximately constant at Mach numbers up to 1.00. At a section normal-force coefficient of 0.32 a shock exists on the upper surface at the lowest test Mach number of 0.79 and supersonic flow exists over approximately 50 percent of the chord on the upper surface. The first indication of a shock on the lower surface occurs at a Mach number of about 0.84. At Mach numbers above 0.95 the shocks on both surfaces occur near the trailing edge and the pressure distrubution over both surfaces is quite similar. An increase in the normal-force coefficient at a Mach number of approximately 0.96 causes a slight increase in the section stability at the higher normal-force coefficients.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 12 December 1956
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    Report Date: September 1950
    No. Pages: 25


  13. FLIGHT MEASUREMENTS WITH THE DOUGLAS D-558-II (BUAERO NO. 37974) RESEARCH AIRPLANE: MEASUREMENTS OF WING LOADS AT MACH NUMBERS UP TO 0.87 , Research Memorandum
    Authors: John P. Mayer, George M. Valentine and Beverly J. Swanson
    Report Number: NACA-RM-L50H16
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Measurements of wing loads at Mach numbers up to 0.87.
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    Report Date: December 1950


  14. TABULATED PRESSURE COEFFICIENTS AND AERODYNAMIC CHARACTERISTICS MEASURED ON THE WING OF THE BELL X-1 AIRPLANE IN LEVEL FLIGHT AT MACH NUMBERS FROM 0.79 TO 1.00 AND IN A PULL-UP AT A MACH NUMBER OF 0.96
    Authors: H. Arthur Carner and Mary M. Payne
    Report Number: NACA-RM-L50H25
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Tabulated pressure coefficients and aerodynamic characteristics are presented for six spanwise stations on the left wing of the Bell X-1 research airplane. The data were obtained in level flight at Mach numbers from 0.79 to 1.00 and in a pull-up to an airplane normal-force coefficient of 0.91 at a Mach number of approximately 0.96.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 12 December 1956
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    Report Date: September 1950
    No. Pages: 43


  15. TABULATED PRESSURE COEFFICIENTS AND AERODYNAMIC CHARACTERISTICS MEASURED ON THE WING OF THE BELL X-1 AIRPLANE IN PULL-UPS AT MACH NUMBERS FROM 0.53 TO 0.99 , Research Memorandum
    Authors: Ronald J. Knapp and Gertrude V. Wilken
    Report Number: NACA-RM-L50H28
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Tabulated pressure coefficients and aerodynamic characteristics are presented for six spanwise stations on the left wing of the Bell X-l research airplane. The data were obtained in 10 pull-ups at Mach numbers from 0.53 to 0.99.
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    Report Date: November 1950


  16. EFFECTS ON THE LATERAL OSCILLATION OF FIXING THE RUDDER AND REFLEXING THE FLAPS ON THE BELL X-1 AIRPLANE
    Authors: Hubert M. Drake
    Report Number: NACA-RM-L50I05
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests have been made on the Bell X-1 airplane having the 10-percent-thick wing and the 8-percent-thick tail to evaluate the effects of fixing the rudder and changing the inclination of the principal axes of inertia by reflexing the landing flaps on the snaking which has been encountered over practically the entire range of Mach number and normal-force coefficient. The data were obtained during power-off glides at altitudes between 32,000 and 16,000 feet. The results showed that fixing the rudder reduced the amplitude of snaking, but did not eliminate it at a Mach numbr of 0.84. It was also found that reflexing the flaps to change the inclination of the principal axis of inertia 1 and 3/4 nose up increased the dynamic lateral stability, but had only a small effect on the snaking oscillation at a Mach number of 0.85.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 19 July 1951
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    Report Date: December 1950
    No. Pages: 13


  17. DETERMINATION OF LONGITUDINAL STABILITY OF THE BELL X-1 AIRPLANE FROM TRANSIENT RESPONSES AT MACH NUMBERS UP TO 1.12 AT LIFT COEFFICIENTS OF 0.3 AND 0.6
    Authors: Ellwyn E. Angle and Euclid C. Holleman
    Report Number: NACA-RM-L50I06A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A number of free-flight transient responses resulting from small stabilizer movements were obtained during flight tests of the Bell X-1 airplane (8-percent-thick wing and 6-percent-thick tail). Responses were analyzed to obtain a measure of the longitudinal stability characteristics of the airplane over the Mach number range from 0.72 to 1.12 at lift coefficients of 0.3 and 0.6. The data presented indicate three significant features: (1) The damping varies greatly with Mach number, maximum damping occurring at Mach numbers of 0.82 and 1.08 and a minimum damping at about 0.93; (2) some uncertainty of damping between Mach numbers of 0.91 to 0.95 appears although good agreement with model tests exists throughout the Mach number range covered; and (3) the static stability of the airplane increases with Mach number to a Mach number of about 0.93 and decreases with further increasing Mach number. Data above a Mach number of 0.90 indicate some lift-coefficient effects. Agreement of the full-scale flight data and model data over the Mach number range is good.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 17 September 1958
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    Report Date: November 1950
    No. Pages: 23


  18. TABULATED PRESSURE COEFFICIENTS AND AERODYNAMIC CHARACTERISTICS MEASURED IN FLIGHT ON THE WING OF THE DOUGLAS D-558-I AIRPLANE FOR A 1 G STALL, A SPEED RUN TO A MACH NUMBER OF 0.90, AND A WIND-UP
    Authors: Earl R. Keener and Mary Pierce
    Report Number: NACA-RM-L50J10
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Tabulated pressure coefficients and aerodynamic characteristics are presented unanalyzed for six spanwise stations on the right wing of the Douglas D-558-I research airplane (BuAero No. 37972). The data were obtained in a 1 g stall at subcritical Mach numbers, in a speed run to a Mach number of 0.90 and in a wind-up turn at a Mach number of 0.86.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 13 September 1954
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    Report Date: December 1950
    No. Pages: 40


  19. FLIGHT MEASUREMENTS WITH THE DOUGLAS D-558-II (BUAERO NO. 37974) RESEARCH AIRPLANE MEASUREMENTS OF THE DISTRIBUTION OF THE AERODYNAMIC LOAD AMONG THE WING, FUSELAGE, AND HORIZONTAL TAIL AT MACH NUMBERS UPS UP TO 0.87
    Authors: John P. Mayer and George M. Valentine
    Report Number: NACA-RM-L50J13
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight measurements of the aerodynamic wing and tail loads have been made on the Douglas D-558-II airplane from which the distribution of the aerodynamic load among the wing, fuselage, and horizontal tail has been determined at Mach numbers up to 0.87. These measurements indicate that, for normal-force coefficients less than 0.7, the distribution of air load among the airplane components does not change appreciably with Mach number at Mach numbers up to 0.87. The measurements also indicate that, for all flight configurations, the increase in airplane normal-force coefficient above the angle of attack at which the wing reaches its maximum normal-force coefficient is due principally to the contribution of the fuselage to the airplane normal-force coefficient.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 24 June 1958
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    Report Date: January 1950
    No. Pages: 31


  20. EFFECTS ON THE SNAKING OSCILLATION OF THE BELL X-1 AIRPLANE OF A TRAILING-EDGE BULB ON THE RUDDER
    Authors: Hubert M. Drake and Harry P. Clagett
    Report Number: NACA-RM-L50K01A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A rudder bulb was installed on the trailing edge of the rudder of the Bell X-1 airplane having the 8-percent-thick wing and 6-percent-thick tail. Several flights were made to investigate the effects of the bulb on the snaking oscillation at Mach numbers between 0.75 and 1.0. It was found that the rudder bulb had no noticeable affect on the snaking oscillation over the Mach number range tested.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 13 September 1954
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    Report Date: January 1950
    No. Pages: 13


  21. TABULATED PRESSURE COEFFICIENTS AND AERODYNAMIC CHARACTERISTICS MEASURED IN FLIGHT ON THE WING OF THE WING OF THE DOUGLAS D-558-I AIRPLANE THROUGHOUT THE NORMAL-FORCE-COEFFICIENT RANGE AT MACH NUMBERS OF 0.67, 0.74, 0.78, AND 0.82
    Authors: Earl R. Keener, James R. Peele and Julia B. Woodbridge
    Report Number: NACA-RM-L50L12A
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Tabulated pressure coefficients and aerodynamic characteristics measured in flight are presented for six spanwise stations on the right wing of the D-558-I research airplane (BuAero No. 37972). The data were obtained throughout the normal-force-coefficient range at Mach numbers of 0.67, 0.74, 0.78, and 0.82. This paper supplements similar tabulated data which have been presented in NACA RM L50J10.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 13 September 1954
    Availability:
        Format(s) on-line:
          Postscript (9,470 KBytes)
          PDF (1,732 KBytes)
    Report Date: January 1950
    No. Pages: 37


  22. RESULTS OBTAINE DURING FLIGHTS 1 TO 6 OF THE NORTHROP X-4 AIRPLANE (A.F. NO. 46-677)
    Authors: James T. Matthews, Jr.
    Report Number: NACA-RM-L9K22
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NACA instruments were installed in the Northrop X-4 number 2 airplane (A.F. No. 46-677) to obtain stability and control data during the acceptance tests conducted by the Northrop Company. The results of flights 1 to 6 are presented in this report. These data were obtained for a center-of-gravity position of about 19.5 to 20.0 percent of the mean aerodynamic chord. The data presented include a time history of a complete pull-up, time histories of several level and accelerated flight runs, and the effect of dive-brake extension on the longitudinal and lateral trim. The pilot reports the mechanical trim device to be unsatisfactory for any stick-free or dynamic stability and control analysis because the stick force cannot be trimmed to zero sufficiently well to permit the stick to be released during a maneuver without the airplane performing a divergence. In addition, the trim device is imoperative when more than 8.0 degrees up elevon angle is required for trim. A short-period longitudinal oscillation with relatively poor damping was present, but this oxcillation was not objectionalble to the pilot. The airplane has a stable variation of longitudinal-control angle with normal-force coefficient for the indicated airspeed ranges of 180 to 300 miles per hour at about 30,000-foot pressure altitude. Extension of the dive brakes up to +- 30 degrees has no appreciable effect on the longitudinal trim at indicated airspeeds of 160 miles per hour with landing gear down, and at airspeeds of 300 miles per hour with the landing gear up, at altitudes of 8,500 and 10,000 feet, respectively. A slight tendency to roll to the left was indicated in the landing-gear-down case.
    Distribution/Availability: Unclassified - Unlimited
    Report declassified: 9 September 1954
    Availability:
        Format(s) on-line:
          Postscript (6,205 KBytes)
          PDF (1,381 KBytes)
    Report Date: January 1950
    No. Pages: 17
 
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