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- 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: Longitudinal pulse maneuvers were conducted on a swept-wing fighter-type airplane for an original-wing and an extended wing-tip configuration at altitudes from 10,000 to 40,000 feet over a Mach number range from 0.36 to 1.45. Variations of the period, damping and the derivatives C (sub) L (sub) sigma, C (sub) m (sub) sigma, and (C sub m sub q + C sub m sub sigma) are presented as functions of Mach number. Comparisons are made with wind-tunnel data. Some consideration is given to pilot opinion in regard to the dynamic longitudinal behavior of the airplane during simulated combat maneuvers.
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Report Date: April 1957
- STATIC-PRESSURE ERROR CALIBRATONS FOR NOSE-BOOM AIRSPEED INSTALLATIONS OF 17 AIRPLANES
Authors: Terry J. Larson, Wendell H. Stillwell and Katharine H. Armistead
Report Number: NACA-RM-H57A02
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A flight investigation was conducted to determine the static-pressure errors for nose-boom airspeed installations of 17 airplanes. The investigation covered both research-type and service-type aircraft. The magnitude of static-pressure errors for the airspeed installations of all the airplanes is shown to vary with airplane geometric characteristics which include nose-boom length, fuselage diameter, and nose fineness ratio. The static-pressure errors for airspeed installations of airplanes with neither extremely blunt nor extremely pointed nose shapes correlate well with the ratio of nose-boom length to effective maximum fuselage diameter. The magnitudes of static-pressure errors vary inversely with this ratio and increase considerably as this ratio decreases below about 1.0.
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Report declassified: 4 February 1959
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Report Date: March 1957
No. Pages: 30
- FLIGHT-DETERMINED STATIC LATERAL STABILITY AND CONTROL CHARACTERISTICS OF A SWEPT-WING FIGHTER AIRPLANE TO A MACH NUMBER OF 1.39.
Authors: Gene J. Matranga and James R. Peele
Report Number: NACA-RM-H57A16
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flight tests were performed with a swept-wing fighter-type airplane at an altitude of 40,000 feet over a Mach number range from 0.72 to 1.39 to determine the lateral stability and control characteristics. Results are presented for three different vertical tails and two different wing areas, and include plots of lateral stability derivatives against Mach number.
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Report Date: March 1957
- FLIGHT-DETERMINED INDUCTION-SYSTEM AND SURGE CHARACTERISTICS OF THE YF-102 AIRPLANE WITH A TWO-SPOOL TURBOJET ENGINE.
Authors: Edwin J. Saltzman
Report Number: NACA-RM-H57C22
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Total-pressure recovery and distortion at the compressor face have been recorded for a twin-side inlet, two-spool turbojet engine combination during turns, sideslips, and speed runs at altitudes between 33,000 and 50,000 feet. In addition, conditions prior to several compressor surges have been recorded. The Mach number range covered extends from about 0.6 to 1.1. The investigation showed that engine surge as experienced is not related to distortion at the compressor face. Mismatching existed for most flight conditions.
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Report Date: June 1957
- AIRPLANE MOTIONS AND LOADS INDUCED BY FLYING THROUGH THE FLOW FIELD GENERATED BY AN AIRPLANE AT LOW SUPERSONIC SPEEDS.
Authors: Gareth H. Jordan, Earl R. Keener and Stanley P. Butchart
Report Number: NACA-RM-H57D17A
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Data are presented for the maximum sideslip angles and vertical-tail loads induced on a swept wing fighter-type airplane as a result of flying through the flow field generated by a similar airplane at low supersonic Mach numbers. These data were obtained during side-by-side passes at various passing rates (5 fps to 50 fps) and interval separation distances. Significant airplane sideslip angles and vertical-tail loads were obtained during close-proximity passes at a passing rate near the natural period of the airplane in yaw.
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Report Date: June 1957
- FLIGHT MEASUREMENTS OF AIRPLANE STRUCTURAL TEMPERATURES AT SUPERSONIC SPEEDS.
Authors: Richard D. Banner
Report Number: NACA-RM-H57D18B
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Skin and structural temperature distributions were obtained during transient supersonic flights of the X-1B and X-1E airplanes at Mach numbers up to approximately 2.0. Extensive temperature measurements were obtained on the X-1B. No critical temperatures were experienced over the range of the test. The measured temperatures were compared with simplified calculations.
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Report Date: June 1957
- FLIGHT MEASUREMENTS AND CALCULATIONS OF WING LOADS AND LOAD DISTRIBUTIONS AT SUBSONIC, TRANSONIC, AND SUPERSONIC SPEEDS.
Authors: Frank S. Malvestuto, Thomas V. Cooney and Earl R. Keener
Report Number: NACA-RM-H57E01
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Presented in this report is a summary of local and net angle-of-attack wing-panel loads measured in flight on six airplanes. In addition, a comparison of these loads measured in flight with calculations based on simple theory is presented.
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Report Date: July 1957
- EFFECT OF WING-MOUNTED EXTERNAL STORES ON THE LIFT AND DRAG OF THE DOUGLAS D-558-II RESEARCH AIRPLANE AT TRANSONIC SPEEDS.
Authors: Jack Nugent
Report Number: NACA-RM-H57E15A
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Lift and drag measurements were made during a flight investigation with the Douglas D-558-II(145) airplane in the basic and 150-gallon DAC store configurations over a Mach number range from 0.48 to 1.03. The addition of stores increased the drag at all Mach numbers tested. Below the drag rise the increase was of about the same magnitude as the increase in wetted area caused by the addition of the stores. The peak lift-drag ratio was reduced by about 14 percent, and for lift coefficients of 0.2 and 0.4 a reduction in drag-rise Mach number was noted. Little change in lift-curve slope was observed for Mach numbers less than about 0.8.
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Report Date: July 1957
- TRANSONIC FLIGHT EVALUATION OF THE EFFECTS OF FUSELAGE EXTENSION AND INDENTATION ON THE DRAG OF A 60 DEGREE DELTA-WING INTERCEPTOR AIRPLANE
Authors: Edwin J. Saltzman and William P. Asher
Report Number: NACA-RM-H57E29
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Lift and drag characteristics of a 60 degree delta-wing interceptor airplane incorporating fuselage extension and indentation have been determined in flight. The data were obtained over the Mach number range from about 0.7 to 1.15 and for altitudes of 25,000, 40,000, and 50,000 feet. These data are compared with the lift and drag characteristics of an airplane with a similar wing, but which did not incorporate modifications to indent or lengthen the fuselage, to determine whether transonic drag was reduced at high Reynolds numbers by improving the cross-sectional-area development. The results of the investigation indicate that anticipated transonic drag reductions have been realized, the reduction amounting to about 0.0050 in drag coefficient at a Mach number of about 1.1. This reduction amounts to about 25 percent of the drag rise for the protorype airplane. The reduced transonic drag of the modified airplane resulted in an improvement in maximum
lift-drag ratio of about 15 percent in the supersonic region. There are significant changes in longitudinal trim which result in less trim drag for the modified airplane. These changes in trim amount to from 1 degree to 2 degrees less control deflection needed by the modified airplane at moderate lift conditions. Three sets of comparable model data are included. These low Reynolds number tests indicated reductions in drag coefficient, due to indenting and extending the fuselage, ranging from 0.0025 to 0.0045 at a Mach numbr of about 1.1.
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Report declassified: 26 September 1957
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Report Date: September 1957
No. Pages: 27
- FLIGHT DATA PERTINENT TO BUFFETING AND MAXIMUM NORMAL-FORCE COEFFICIENT OF THE DOUGLAS X-3 RESEARCH AIRPLANE.
Authors: Thomas F. Baker, James A. Martin and Betty J. Scott
Report Number: NACA-RM-H57H09
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The X-3 airplane, which has a straight, 4.5-percent-thick wing, was flown to maximum wing lift at Mach numbers from 0.7 to 1.1 at an average altitude of 30,000 feet. Airplane and wing maximum normal-force coefficients and buffeting characteristics were determined. Wing maximum normal-force coefficients at low supersonic speeds were almost twice the values at subsonic speed (M nearly equal to 0.8). At transonic speeds the buffet boundary abruptly increased, rather than decreased, with Mach number. The buffeting encountered did not constitute either an operational or a structural problem. The effective longitudinal maneuverability limit was defined by maximum wing lift. Limited data at subsonic speeds on the effects on lift and buffeting of deflecting the wing leading edge flaps 7 degrees are included.
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Report Date: November 1957
- EFFECTS OF WING-MOUNTED EXTERNAL STORES ON THE LONGITUDINAL AND LATERAL HANDLING QUALITIES OF THE DOUGLAS D-558-II RESEARCH AIRPLANE.
Authors: Jack Fischel, Robert W. Darville and Donald Reisert
Report Number: NACA-RM-H57H12
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The subsonic and transonic handling qualities of the Douglas D-558-II research airplane were investigated with several configuration of midsemispan external stores in the altitude region between 20,000 and 40,000 feet. The configurations tested consisted of an underslung pylon on each wing, pylons plus simulated DAC (Douglas Aircraft Co.) 1,000-pound bombs, and pylons plus DAC 150-gallon-fuel tanks. Comparisons of the results obtained were made with comparable data from the clean airplane. The trends exhibited in the characteristics measured with each configuration were generally the same as for the clean airplane; however, significant changes in the magnitude of the parameters measured with the pylon-tank configuration were sometimes apparent, particularly at the higher speeds tested.
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Report Date: October 1957
- FLIGHT INVESTIGATION OF THE TRANSONIC LONGITUDINAL AND LATERAL HANDLING QUALITIES OF THE DOUGLAS X-3 RESEARCH AIRPLANE.
Authors: Jack Fischel , Euclid C. Holleman and Robert A. Tremant
Report Number: NACA-RM-H57I05
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A flight investigation was performed to determine the longitudinal and lateral handling qualities of the Douglas X-3 research airplane in the clean configuration and with wing leading-edge flaps deflected. Static and dynamic stability and control characteristics were determined during trimmed and maneuvering flight at an average altitude of 30,000 feet over a Mach number range from 0.7 to 1.16. Statically and dynamically determined stability and control derivatives are presented, as well as pilot evaluation of the airplane.
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Report Date: December 1957
- IN-FLIGHT GAINS REALIZED BY MODIFYING A TWIN SIDE-INLET INDUCTION SYSTEM.
Authors: Edwin J. Saltzman
Report Number: NACA-RM-H57J09
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The effects of modifying a twin side-inlet duct system have been recorded and analyzed over an altitude range from about 25,000 to 51,000 feet and throughout the transonic region to a Mach number of about 1.2. The modification consisted primarily of redesigning the inlet lip, increasing the cross-sectional area of the inlet and diffuser, and adding a region of duct contraction ahead of the engine. These changes greatly improved the pressure-recovery characteristics and provided a 50-percent reduction in compressor-face distortion (pressure-profile variation).
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Report Date: December 1957
- AN ANALYSIS OF VERTICAL-TAIL LOADS MEASURED IN FLIGHT ON A SWEPT-WING BOMBER AIRPLANE.
Authors: William A. McGowan and T. V. Cooney
Report Number: NACA-RM-L57B19
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An analysis is presented of vertical-tail loads measured on a swept-wing bomber airplane at altitudes to 35,000 feet and Mach numbers to 0.82. Flight data obtained from rudder-step, rudder-pulse, aileronroll, and steady-sideslip maneuvers were used in the analysis to determine lift-curve slopes, centers of pressure, and wing-fuselage, tail, and airplane static-directional-stability parameters. Results are compared, where possible, with values used in design and with theoretical values. Theoretical values of the lift-curve slopes were in agreement with flight values when fuselage flexibility was considered.
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Report Date: May 1957
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