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  1. INVESTIGATION OF COAXIAL JET NOISE AND INLET CHOKING USING AN F-111A AIRPLANE , Technical Note
    Authors: T. W. Putnam
    Report Number: NASA-TN-D-7376
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Measurements of engine noise generated by an F-111A airplane positioned on a thrustmeasuring platform were made at angles of 0 deg to 160 deg from the aircraft heading. Sound power levels, power spectra, and directivity patterns are presented for jet exit velocities between 260 feet per second and 2400 feet per second. The test results indicate that the total acoustic power was proportional to the eighth power of the core jet velocity for core exhaust velocities greater than 300 meters per second (985 feet per second) and that little or no mixing of the core and fan streams occurred. The maximum sideline noise was most accurately predicted by using the average jet velocity for velocities above 300 meters per second (985 feet per second). The acoustic power spectrum was essentially the same for the single jet flow of afterburner operation and the coaxial flow of the nonafterburning condition. By varying the inlet geometry and cowl position, reductions in the sound pressure level of the blade passing frequency on the order of 15 decibels to 25 decibels were observed for inlet Mach numbers of 0.8 to 0.9.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 71
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    Report Date: August 1973
    No. Pages: 34
    Keywords:      Aircraft noise; F-111 aircraft; Jet exhaust; Noise intensity; Noise spectra.


  2. WIND-TUNNEL CALIBRATION AND REQUIREMENTS FOR IN-FLIGHT USE OF FIXED HEMISPHERICAL HEAD ANGLE-OF-ATTACK AND ANGLE-OF-SIDESLIP SENSORS , Technical Note
    Authors: E. J. Montoya
    Report Number: NASA-TN-D-6986
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Wind-tunnel tests were conducted with three different fixed pressure-measuring hemispherical head sensor configurations which were strut-mounted on a nose boom. The tests were performed at free-stream Mach numbers from 0.2 to 3.6. The boom-angle-of-attack range was -6 to 15 deg, and the angle-of-sideslip range was -6 to 6 deg. The test Reynolds numbers were from 3.28 million to 65.6 million per meter. The results were used to obtain angle-of-attack and angle-of-sideslip calibration curves for the configurations. Signal outputs from the hemispherical head sensor had to be specially processed to obtain accurate real-time angle-of-attack and angle-of-sideslip measurements for pilot displays or aircraft systems. Use of the fixed sensors in flight showed them to be rugged and reliable and suitable for use in a high temperature environment.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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    Report Date: March 1973
    No. Pages: 47
    Keywords:      Angle of attack; Free flight; Sensors; Sideslip; Wind tunnel models.


  3. ANALYTICAL STUDY OF TAKEOFF AND LANDING PERFORMANCE FOR A JET STOL TRANSPORT CONFIGURATION WITH FULL-SPAN, EXTERNALLY BLOWN, TRIPLE-SLOTTED FLAPS , Technical Note
    Authors: H. P. Washington and J. T. Gibbons
    Report Number: NASA-TN-D-7441
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Takeoff and landing performance characteristics and field length requirements were determined analytically for a jet STOL transport configuration with full-span, externally blown, tripleslotted flaps. The configuration had a high wing, high T-tail, and four pod-mounted high-bypass-ratio turbofan engines located under and forward of the wing. One takeoff and three approach and landing flap settings were evaluated. The effects of wing loading, thrust-to-weight ratio, weight, ambient temperature, altitude on takeoff and landing field length requirements are discussed.
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    Subject Category: 05
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    Report Date: October 1973
    No. Pages: 49
    Keywords:      Aerodynamic characteristics; Aircraft performance; Externally blown flaps; Short takeoff aircraft; Transport aircraft.


  4. LOCAL SKIN FRICTION COEFFICIENTS AND BOUNDARY LAYER PROFILES OBTAINED IN FLIGHT FROM THE XB-70-1 AIRPLANE AT MACH NUMBERS UP TO 2.5 , Technical Note
    Authors: D. F. Fisher and E. J. Saltzman
    Report Number: NASA-TN-D-7220
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Boundary-layer and local friction data for Mach numbers up to 2.5 and Reynolds numbers up to 3.6 x 10 to the 8th power were obtained in flight at three locations on the XB-70-1 airplane: the lower forward fuselage centerline (nose), the upper rear fuselage centerline, and the upper surface of the right wing. Local skin friction coefficients were derived at each location by using (1) a skin friction force balance, (2) a Preston probe, and (3) an adaptation of Clauser's method which derives skin friction from the rake velocity profile. These three techniques provided consistent results that agreed well with the von Karman-Schoenherr relationship for flow conditions that are quasi-two-dimensional. At the lower angles of attack, the nose-boom and flow-direction vanes are believed to have caused the momentum thickness at the nose to be larger than at the higher angles of attack. The boundary-layer data and local skin friction coefficients are tabulated. The wind-tunnel-model surface-pressure distribution ahead of the three locations and the flight surface-pressure distribution ahead of the wing location are included.
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    Subject Category: 34
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    Report Date: June 1973
    No. Pages: 70
    Keywords:      Aerodynamic characteristics; Aerodynamic coefficients; B-70 aircraft; Boundary layer flow; Skin friction.


  5. FLIGHT INVESTIGATION OF XB-70 STRUCTURAL RESPONSE TO OSCILLATORY AERODYNAMIC SHAKER EXCITATION AND CORRELATION WITH ANALYTICAL RESULTS , Technical Note
    Authors: J. M. McKay, E. E. Kordes and J. H. Wykes
    Report Number: NASA-TN-D-7227
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The low frequency symmetric structural response and damping characteristics of the XB-70 airplane were measured at four flight conditions: heavyweight at a Mach number of 0.87 at an altitude of 7620 meters (25,000 feet); lightweight at a Mach number of 0.86 at an altitude of 7620 meters (25,000 feet); a Mach number of 1.59 at an altitude of 11,918 meters (39.100 feet); and a Mach number of 2.38 and an altitude of 18,898 meters (62,000 feet). The flight data are compared with the response calculated by using early XB-70 design data and with the response calculated with mass, structural, and aerodynamic data updated to reflect as closely as possible the airplane characteristics at three of the flight conditions actually flown.
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    Subject Category: 39
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    Report Date: April 1973
    No. Pages: 124
    Keywords:      B-70 aircraft; Dynamic response; Flight tests; Structural stability; Vibration damping.


  6. MEASUREMENTS OF SURFACE-PRESSURE FLUCTUATIONS ON THE XB-70 AIRPLANE AT LOCAL MACH NUMBERS UP TO 2.45 , Technical Note
    Authors: T. L. Lewis, J. B. Dods, Jr. and R. D. Hanly
    Report Number: NASA-TN-D-7226
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Measurements of surface-pressure fluctuations were made at two locations on the XB-70 airplane for nine flight-test conditions encompassing a local Mach number range from 0.35 to 2.45. These measurements are presented in the form of estimated power spectral densities, coherence functions, and narrow-band-convection velocities. The estimated power spectral densities compared favorably with wind-tunnel data obtained by other experimenters. The coherence function and convection velocity data supported conclusions by other experimenters that low-frequency surface-pressure fluctuations consist of small-scale turbulence components with low convection velocity.
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    Subject Category: 02
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    Report Date: March 1973
    No. Pages: 35
    Keywords:      Aerodynamic characteristics; Aerodynamic loads; B-70 aircraft; Pressure distribution; Pressure measurement.


  7. LOCAL FLOW MEASUREMENTS AT THE INLET SPIKE TIP OF A MACH 3 SUPERSONIC CRUISE AIRPLANE , Technical Note
    Authors: H. J. Johnson and E. J. Montoya
    Report Number: NASA-TN-D-6987
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The flow field at the left inlet spike tip of a YF-12A airplane was examined using at 26 deg included angle conical flow sensor to obtain measurements at free-stream Mach numbers from 1.6 to 3.0. Local flow angularity, Mach number, impact pressure, and mass flow were determined and compared with free-stream values. Local flow changes occurred at the same time as free-stream changes. The local flow usually approached the spike centerline from the upper outboard side because of spike cant and toe-in. Free-stream Mach number influenced the local flow angularity; as Mach number increased above 2.2, local angle of attack increased and local sideslip angle decreased. Local Mach number was generally 3 percent less than free-stream Mach number. Impact-pressure ratio and mass flow ratio increased as free-stream Mach number increased above 2.2, indicating a beneficial forebody compression effect. No degradation of the spike tip instrumentation was observed after more than 40 flights in the high-speed thermal environment encountered by the airplane. The sensor is rugged, simple, and sensitive to small flow changes. It can provide accurate imputs necessary to control an inlet.
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    Subject Category: 02
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    Report Date: May 1973
    No. Pages: 42
    Keywords:      Flow measurement; Impact loads; Inlet flow; Mach number; Supersonic aircraft.


  8. MODELING OF AIRPLANE PERFORMANCE FROM FLIGHT-TEST RESULTS AND VALIDATION WITH AN F-104G AIRPLANE
    Authors: R. T. Marshall and W. G Schweikhard
    Report Number: NASA-TN-D-7137
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A technique of defining an accurate performance model of an airplane from limited flight-test data and predicted aerodynamic and propulsion system characteristics is developed. With the modeling technique, flight-test data from level accelerations are used to define a 1g performance model for the entire flight envelope of an F-104G airplane. The performance model is defined in terms of the thrust and drag of the airplane and can be varied with changes in ambient temperature or airplane weight. The model predicts the performance of the airplane within 5 percent of the measured flight-test data. The modeling technique could substantially reduce the time required for performance flight testing and produce a clear definition of the thrust and drag characteristics of an airplane.
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    Subject Category: 05
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    Report Date: February 1973
    No. Pages: 30
    Keywords:      Aerodynamic characteristics; Aerodynamic coefficients; Aircraft performance; Performance prediction; Propulsion system performance.


  9. PRESSURES AND TEMPERATURES ON THE LOWER SURFACES OF AN EXTERNALLY BLOWN FLAP SYSTEM DURING FULL-SCALE GROUND TESTS , Technical Note
    Authors: D. L. Hughes
    Report Number: NASA-TN-D-7138
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Full-scale ground tests of an externally blown flap system were made using the wing of an F-111B airplane and a CF700 engine. Pressure and temperature distributions were determined on the undersurface of the wing, vane, and flap for two engine exhaust nozzles (conical and daisy) at several engine power levels and engine/wing positions. The test were made with no airflow over the wing. The wing sweep angle was fixed at e6 deg; and the angle of incidence between the engine and the wing was fixed at 3 deg; and the flap was in the retracted, deflected 35 deg, and deflected 60 deg positions. The pressure load obtained by integrating the local pressures on the undersurface of the flap, F sub p was approximately three times greater at the 60 deg flap position than at the 35 deg flap position. At the 60 deg flap position, F sub p was between 40 percent and 55 percent of the engine thrust over the measured range of thrust. More than 90 percent of F sub p was contained within plus or minus 20 percent of the flap span centered around the engine exhaust centerline with both nozzle configurations. Maximum temperatures recorded on the flaps were 218 C (424 F) and 180 C (356 F) for the conical and daisy nozzles, repectively.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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    Report Date: January 1973
    No. Pages: 34
    Keywords:      CF-700 engine; Externally blown flaps; F-111 aircraft; Pressure distribution; Temperature distribution.


  10. FLIGHT INVESTIGATION OF A STRUCTURAL MODE CONTROL SYSTEM FOR THE XB-70 AIRCRAFT , Technical Note
    Authors: W. P. Lock, E. E. Kordes, J. M. McKay and J. H. Wykes
    Report Number: NASA-TN-D-7420
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight investigation of a structural mode control system termed identical location of accelerometer and force (ILAF) was conducted on the XB-70-1 airplane. During the first flight tests, the ILAF system encountered localized structural vibration problems requiring a revision of the compensating network. After modification, successful structural mode control that did not adversely affect the rigid body dynamics was demonstrated. The ILAF system was generally more effective in supersonic than subsonic flight, because the conditions for which the system was designed were more nearly satisfied at supersonic speeds. The results of a turbulence encounter at a Mach number of 1.20 and an altitude of 9754 meters indicated that the ILAF system was effective in reducing the vehicle's response at this flight condition. An analytical study showed that the addition of a small canard to the modal suppression system would greatly improve the automatic control of the higher frequency symmetric modes.
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    Subject Category: 05
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    Report Date: October 1973
    No. Pages: 83
    Keywords:      Aircraft equipment; Aircraft performance; Aircraft stability; B-70 aircraft; Flight control.


  11. DESIGN AND FLIGHT TESTING OF A NULLABLE COMPRESSOR FACERAKE
    Authors: J. K. Holzman and G. A. Payne
    Report Number: NASA-TN-D-7162
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A compressor face rake with an internal valve arrangement to permit nulling was designed, constructed, and tested in the laboratory and in flight at the NASA Flight Research Center. When actuated by the pilot in flight, the nullable rake allowed the transducer zero shifts to be determined and then subsequently removed during data reduction. Design details, the fabrication technique, the principle of operation, brief descriptions of associated digital zero-correction programs and the qualification tests, and test results are included. Sample flight data show that the zero shifts were large and unpredictable but could be measured in flight with the rake. The rake functioned reliably and as expected during 25 hours of operation under flight environmental conditions and temperatures from 230 K (-46 F) to greater than 430 K (314 F). The rake was nulled approximately 1000 times. The in-flight zero-shift measurement technique, as well as the rake design, was successful and should be useful in future applications, particularly where accurate measurements of both steady-state and dynamic pressures are required under adverse environmental conditions.
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    Subject Category: 34
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    Report Date: January 1973
    No. Pages: 17
    Keywords:      Equipment specifications; In-flight monitoring; Jet aircraft; Pressure sensors; Reliability engineering.


  12. GROUND AND FLIGHT EXPERIENCE WITH A STRAPDOWN INERTIAL MEASURING UNIT AND A GENERAL PURPOSE AIRBORNE DIGITAL COMPUTER
    Authors: T. D. Wolf and R. C. McCracken
    Report Number: NASA-TM-X-2848
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Ground and flight tests were conducted to investigate the problems associated with using a strapdown inertial flight data system. The objectives of this investigation were to develop a three axis inertial attitude reference system, to evaluate a self-alignment technique, and to examine the problem of time-sharing a general purpose computer for the several tasks required of it. The performance of the strapdown platform/computer system that was developed was sufficiently accurate for the tasks attempted. For flights on the order of 45 minutes duration, attitude angle errors of + or -.035 radian (+ or -2 deg) in all axes were observed. Laboratory tests of the self-alignment technique gave accuracies of + or -.00075 radian in pitch and roll axes and + or -0.0045 radian in the yaw axis. Self-alignment flight results were inconsistent, since a stable solution was not obtained on windy days because of aircraft rocking motions.
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    Subject Category: 04
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    Report Date: August 1973
    No. Pages: 34
    Keywords:      Airborne/spaceborne computers; Attitude indicators; Computer techniques; Performance tests; Strapdown inertial guidance.


  13. RESULTS OF GROUND VIBRATION TESTS ON A YF-12 AIRPLANE
    Authors: R. J. Wilson, F. W. Cazier, Jr. and R. R. Larson
    Report Number: NASA-TM-X-2880
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Ground vibration tests were conducted on a YF-12 airplane. To approximate a structural free-free boundary condition during the tests, each of the landing gears was supported on a support system designed to have a low natural frequency. The test equipment and the procedures used for the ground vibration tests are described. The results are presented in the form of frequency response data, measured mode lines, and elastic mode shapes for the wing/body, rudder, and fuselage ventral fin. In the frequency range between 3.4 cps and 28.8 cps, nine symmetrical wing/body modes, six antisymmetrical wing/body modes, two rudder modes, and one ventral fin mode were measured.
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    Subject Category: 39
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    Report Date: August 1973
    No. Pages: 94
    Keywords:      Airframes; Ground tests; Military aircraft; Reconnaissance aircraft; Structural analysis.


  14. PRELIMINARY FLIGHT EVALUATION OF A PAINTED DIAMOND ON A RUNWAY FOR VISUAL INDICATION OF GLIDE SLOPE
    Authors: S. W. Gee and R. C. McCracken
    Report Number: NASA-TM-X-2849
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A diamond sized to appear equidimensional when viewed from a 3.6 deg slide slope was painted on the end of a small general aviation airport runway, and a series of flights was made to evaluate its usage as a piloting aid. The pilots could detect and fly reasonably close to the glide slope projected by the diamond. The flight path oscillations that were recorded during approaches using the diamond were not significantly different from the oscillations that were recorded without the diamond; the difference that did exist could be attributed to converging on a known projected glide slope in one case, and flying an unknown, random glide slope in the other. The results indicated that the diamond would be effective as a means of intercepting and controlling a predetermined glide slope. Other advantages of the diamond were positive runway identification and greater aim point visibility. The major disadvantage was a tendency to overconcentrate on the diamond and consequently to neglect cockpit instruments and airport traffic.
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    Subject Category: 05
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    Report Date: August 1973
    No. Pages: 22
    Keywords:      Aircraft landing; Approach control; Flight paths; Flight safety; Visual perception.


  15. FLIGHT-MEASURED BASE PRESSURE COEFFICIENTS FOR THICKBOUNDARY-LAYER FLOW OVER AN AFT-FACING STEP FOR MACH NUMBERS FROM 0.4TO 2.5
    Authors: S. A. Goecke
    Report Number: NASA-TN-D-7202
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A 0.56-inch thick aft-facing step was located 52.1 feet from the leading edge of the left wing of an XB-70 airplane. A boundary-layer rake at a mirror location on the right wing was used to obtain local flow properties. Reynolds numbers were near 10 to the 8th power, resulting in a relatively thick boundary-layer. The momentum thickness ranged from slightly thinner to slightly thicker than the step height. Surface static pressures forward of the step were obtained for Mach numbers near 0.9, 1.5, 2.0, and 2.4. The data were compared with thin boundary-layer results from flight and wind-tunnel experiments and semiempirical relationships. Significant differences were found between the thick and the thin boundary-layer data.
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    Subject Category: 34
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    Report Date: May 1973
    No. Pages: 33
    Keywords:      B-70 aircraft; Base pressure; Boundary layer flow; Discontinuity; Flight tests.


  16. STEADY-STATE AND DYNAMIC PRESSURE PHENOMENA IN THE PROPULSION SYSTEM OF AN F-111A AIRPLANE
    Authors: F. W. Burcham, D. L. Hughes and J. K. Holzman
    Report Number: NASA-TN-D-7328
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests were conducted with two F-111A airplanes to study the effects of steady-state and dynamic pressure phenomena on the propulsion system. Analysis of over 100 engine compressor stalls revealed that the stalls were caused by high levels of instantaneous distortion. In 73 percent of these stalls, the instantaneous circumferential distortion parameter, k sub theta, exhibited a peak just prior to stall higher than any previous peak. The K sub theta parameter was a better indicator of stall than the distortion factor, k sub d, and the maximum-minus-minimum distortion parameter, d, was poor indicator of stall. Inlet duct resonance occurred in both F-111A airplanes and is believed to have been caused by oscillations of the normal shock wave from an internal to an external position. The inlet performance of the two airplanes was similar in terms of pressure recovery, distortion, and turbulence, and there was good agreement between flight and wind-tunnel data up to a Mach number of approximately 1.8.
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    Subject Category: 07
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    Report Date: July 1973
    No. Pages: 102
    Keywords:      Aerodynamic stalling; Air flow; Aircraft engines; Engine failure; F-111 aircraft.


  17. EFFECTS OF FLAPS ON BUFFET CHARACTERISTICS AND WIND-ROCK ONSET OF AN F-8C AIRPLANE AT SUBSONIC AND TRANSONIC SPEEDS
    Authors: R. C. Monaghan and E. L. Friend
    Report Number: NASA-TM-X-2873
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Wind-up-turn maneuvers were performed to establish the values of airplane normal force coefficient for buffet onset, wing-rock onset, and buffet loads with various combinations of leading- and trailing-edge flap deflections. Data were gathered at both subsonic and transonic speeds covering a range from Mach 0.64 to Mach 0.92. Buffet onset and buffet loads were obtained from wingtip acceleration and wing-root bending-moment data, and wing-rock onset was obtained from airplane roll rate data. Buffet onset, wing-rock onset, and buffet loads were similarly affected by the various combinations of leading- and training-edge flaps. Subsonically, the 12 deg leading-edge-flap and trailing-edge-flap combination was most effective in delaying buffet onset, wing-rock onset, and equivalent values of buffet loads to a higher value of airplane normal force coefficient. This was the maximum flap deflection investigated. Transonically, however, the optimum leading-edge flap position was generally less than 12 deg.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05
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    Report Date: August 1973
    No. Pages: 29
    Keywords:      Aircraft performance; Aircraft stability; Buffeting; F-8 aircraft; Flight tests.


  18. FLIGHT-MEASURED X-24A LIFTING BODY CONTROL SURFACE HINGE MOMENTS AND CORRELATION WITH WIND TUNNEL PREDICTIONS
    Authors: M. H. Tang and G. P. E. Pearson
    Report Number: NASA-TM-X-2816
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Control-surface hinge-moment measurements obtained in the X-24A lifting body flight-test program are compared with results from wind-tunnel tests. The effects of variations in angle of attack, angle of sideslip, rudder bias, rudder deflection, upper-flap deflection, lower-flap deflection, Mach number, and rocket-engine operation on the control-surface hinge moments are presented. In-flight motion pictures of tufts attached to the inboard side of the right fin and the rudder and upper-flap surfaces are discussed.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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    Report Date: June 1973
    No. Pages: 39
    Keywords:      Aerodynamic forces; Control surfaces; Flight tests; Wind tunnel tests; X-24 aircraft.


  19. DEVELOPMENT AND FLIGHT TEST OF A DIGITAL FLY-BY-WIRE F8.
    Authors: B. A. Peterson, G. E. Krier and C. R. Jarvis
    Report Number: H-750
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The objective of the F-8 digital fly-by-wire program is to establish a technology base for the implementation of advanced flight control systems. The central element is the Apollo Lunar Guidance Computer (LGC). This versatile computer ran over 2000 hours in support of fly-by-wire without a failure. Difficulties encountered in the first flights were corrected rapidly and simply by changes in the erasable software memory. Control-configured vehicles offer significant weight-saving possibilities.
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    Subject Category: 05
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    Report Date: January 1973
    No. Pages: 15
    Keywords:      Acceleration (physics); Digital command systems; F-8 aircraft; Flight tests; Fly by wire control.
    Notes: Society of Experimental Test Pilots, Technical Review, vol. 11, no. 2, 1973, p. 57-71.


  20. EXPLICIT DETERMINATION OF LATERAL-DIRECTIONAL STABILITY AND CONTROL DERIVATIVES BY SIMUTANEOUS TIME VECTOR ANALYSIS OF TWO MANEUVERS
    Authors: G. B. Gilyard
    Report Number: NASA-TM-X-2722
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An extension of the time vector technique for determining stability and control derivatives from flight data is formulated. The technique provides for explicit determination of derivatives by means of simultaneous analysis of two maneuvers which differ by a dependent control input. The control derivatives for the dependent input are also explicitly determined. This extended technique is preferable to the application of the time vector method to single maneuvers in that no estimates of derivatives are required. An example illustrating the application of the technique is given.
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    Subject Category: 05
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    Report Date: February 1973
    No. Pages: 12
    Keywords:      Aerodynamic characteristics; Aircraft performance; Flight control; Stability derivatives; Vector analysis.


  21. EXPERIMENTS TO STUDY STRAIN GAGE LOAD CALIBRATIONS ON A WING STRUCTURE AT ELEVATED TEMPERATURES
    Authors: R. C. Monaghan and R. A. Fields
    Report Number: NASA-TN-D-7390
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Laboratory experiments were performed to study changes in strain-gage bridge load calibrations on a wing structure heated to temperatures of 200 F, 400 F, and 600 F. Data were also obtained to define the experimental repeatability of strain-gage bridge outputs. Experiments were conducted to establish the validity of the superposition of bridge outputs due to thermal and mechanical loads during a heating simulation of Mach 3 flight. The strain-gage bridge outputs due to load cycle at each of the above temperature levels were very repeatable. A number of bridge calibrations were found to change significantly as a function of temperature. The sum of strain-gage bridge outputs due to individually applied thermal and mechanical loads compared well with that due to combined or superimposed loads. The validity of superposition was, therefore, established.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 39
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    Report Date: August 1973
    No. Pages: 27
    Keywords:      Aerodynamic heating; Aerodynamic loads; Aircraft structures; Strain gages; Wings.


  22. APPLICATION OF ADVANCED CONTROL SYSTEM AND DISPLAY TECHNOLOGYTO GENERAL AVIATION.
    Authors: M. R. Barber
    Report Number: H-768
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: No abstract available.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05
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    Report Date: April 1973
    No. Pages: 12
    Keywords:      Aircraft control; Attitude control; Automatic pilots; Display devices; General aviation aircraft.
    Notes: Society of Automotive Engineers, Business Aircraft Meeting, Wichita, Kan., Apr. 3-6, 1973. SAE Paper 730321.


  23. DEVELOPMENT OF A LOW-COST FLIGHT DIRECTOR SYSTEM FOR GENERALAVIATION.
    Authors: S. W. Gee and N. A. Servais
    Report Number: H-769
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: No abstract available.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 35
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    Report Date: April 1973
    No. Pages: 10
    Keywords:      Airborne/spaceborne computers; Aircraft control; Cost reduction; Display devices; Flight control.
    Notes: Society of Automotive Engineers, Business Aircraft Meeting, Wichita, Kan., Apr. 3-6, 1973. SAE Paper 730331.


  24. SEPARATE SURFACES FOR AUTOMATIC FLIGHT CONTROLS.
    Authors: J. Roskam, M. R. Barber and P. C. Loschke
    Report Number: H-770
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The purpose of this paper is to describe an investigation of separate surface stability augmentation systems for general aviation aircraft. The program objective were twofold: first, a wind tunnel program to determine control effectiveness of separate surfaces in the presence of main surfaces, and hinge moment feedback from separate surfaces via the main surfaces to the pilot; second, a theoretical study to determine the minimum performance of actuators and sensors that can be tolerated, the best slaving gains to be used with separate surfaces, and control authority needed for proper operation under direct pilot control, under autopilot control, and in failure situations. On the basis of the results obtained, it has been concluded that separate surface systems are feasible and advantageous for use in general aviation aircraft.
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    Report Date: April 1973
    No. Pages: 13
    Keywords:      Automatic flight control; Civil aviation; Feedback control; Flight stability tests; Wind tunnel stability tests.
    Notes: Society of Automotive Engineers, Business Aircraft Meeting, Wichita, Kan., Apr. 3-6, 1973. SAE Paper 73034.


  25. REVIEW OF DRAG MEASUREMENTS FROM FLIGHT TESTS OF MANNEDAIRCRAFT WITH COMPARISONS TO WIND-TUNNEL PREDICTIONS
    Authors: J. S. Pyle and E. J. Saltzman
    Report Number: H-772
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In-flight studies of the overall and local components of drag of many types of aircraft were conducted. The primary goal of these studies was to evaluate wind-tunnel and semiempirical prediction methods. Some evaluations are presented in this paper which may be summarized by the following observations: Wind-tunnel predictions of overall vehicle drag can be accurately extrapolated to flight Reynolds numbers, provided that the base drag is removed and the boattail areas on the vehicle are small. The addition of ablated roughness to lifting body configurations causes larger losses in performance and stability than would be expected from the added friction drag due to the roughness. Successful measurements of skin friction have been made in flight to Mach numbers above 4. A reliable inflatable deceleration device was demonstrated in flight which effectively stabilizes and decelerates a lifting aircraft at supersonic speeds.
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    Subject Category: 02
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    Report Date: October 1973
    No. Pages: 12
    Keywords:      Aerodynamic drag; Aircraft configurations; Drag measurement; Flight characteristics; Flight tests.
    Notes: In AGARD Aerodyn. Drag. AGARD CP-124, Paper 26.


  26. COMPARISONS OF PREDICTIONS OF THE XB-70-1 LONGITUDINAL STABILITY AND CONTROL DERIVATIVES WITH FLIGHT RESULTS FOR SIX FLIGHT CONDITIONS
    Authors: C. H. Wolowicz and R. B. Yancey
    Report Number: NASA-TM-X-2881
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Preliminary correlations of flight-determined and predicted stability and control characteristics of the XB-70-1 reported in NASA TN D-4578 were subject to uncertainties in several areas which necessitated a review of prediction techniques particularly for the longitudinal characteristics. Reevaluation and updating of the original predictions, including aeroelastic corrections, for six specific flight-test conditions resulted in improved correlations of static pitch stability with flight data. The original predictions for the pitch-damping derivative, on the other hand, showed better correlation with flight data than the updated predictions. It appears that additional study is required in the application of aeroelastic corrections to rigid model wind-tunnel data and the theoretical determination of dynamic derivatives for this class of aircraft.
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    Report Date: August 1973
    No. Pages: 44
    Keywords:      Aircraft control; Aircraft performance; Aircraft stability; B-70 aircraft; Data acquisition.


  27. AIRFRAME/PROPULSION SYSTEM INTERACTIONS - AN IMPORTANT FACTORIN SUPERSONIC AIRCRAFT FLIGHT CONTROL.
    Authors: D. T. Berry and G. B. Gilyard
    Report Number: H-774
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The demands of supersonic flight have resulted in propulsion system features that have a significant influence on aircraft flight control. Data from a Mach 3 cruise airplane show that airframe/propulsion system interactions can reduce phugoid and dutch-roll damping, increase vehicle sensitivity to atmospheric disturbances, alter the effective static and dynamic stability of the aircraft, and produce moments as strong as aerodynamic controls. In turn, these effects can lead to large aircraft excursions or high pilot workload, or both, and place increased demands on stability augmentation systems and aerodynamic controls. A need to integrate flight control and propulsion control in advanced vehicles is indicated.
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    Report Date: August 1973
    No. Pages: 8
    Keywords:      Aircraft engines; Airframes; Engine control; Flight control; Propulsion system performance.
    Notes: American Institute of Aeronautics and Astronautics, Guidance and Control Conference, Key Biscayne, Fla., Aug. 20-22, 1973. AIAA Paper 73-831.


  28. STUDY OF CONTROL SYSTEM EFFECTIVENESS IN ALLEVIATING VORTEX WAKEUPSETS.
    Authors: W. A. Johnson and H. A. Rediess
    Report Number: H-777
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The problem of an airplane being upset by encountering the vortex wake of a large transport on takeoff or landing is currently receiving considerable attention. This paper describes the technique and results of a study to assess the effectiveness of automatic control systems in alleviating vortex wake upsets. A six-degree-of-freedom nonlinear digital simulation was used for this purpose. The analysis included establishing the disturbance input due to penetrating a vortex wake from an arbitrary position and angle. Simulations were computed for both a general aviation airplane and a commercial jet transport. Dynamic responses were obtained for the penetrating aircraft with no augmentation and with various command augmentation systems. The results of this preliminary study indicate that it is feasible to use an automatic control system to alleviate vortex encounter upsets.
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    Subject Category: 02
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    Report Date: August 1973
    No. Pages: 10
    Keywords:      Aircraft control; Aircraft wakes; Automatic flight control; Digital simulation; System effectiveness.
    Notes: American Institute of Aeronautics and Astronautics, Guidance and Control Conference, Key Biscayne, Fla., Aug. 20-22, 1973. AIAA Paper 73-833.


  29. FLIGHT CALIBRATION TESTS OF A NOSE-BOOM-MOUNTED FIXED HEMISPHERICAL FLOW-DIRECTION SENSOR
    Authors: K. H. Armistead and L. D. Webb
    Report Number: NASA-TN-D-7461
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight calibrations of a fixed hemispherical flow angle-of-attack and angle-of-sideslip sensor were made from Mach numbers of 0.5 to 1.8. Maneuvers were performed by an F-104 airplane at selected altitudes to compare the measurement of flow angle of attack from the fixed hemispherical sensor with that from a standard angle-of-attack vane. The hemispherical flow-direction sensor measured differential pressure at two angle-of-attack ports and two angle-of-sideslip ports in diametrically opposed positions. Stagnation pressure was measured at a center port. The results of these tests showed that the calibration curves for the hemispherical flow-direction sensor were linear for angles of attack up to 13 deg. The overall uncertainty in determining angle of attack from these curves was plus or minus 0.35 deg or less. A Mach number position error calibration curve was also obtained for the hemispherical flow-direction sensor. The hemispherical flow-direction sensor exhibited a much larger position error than a standard uncompensated pitot-static probe.
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    Report Date: October 1973
    No. Pages: 29
    Keywords:      Aircraft instruments; F-104 aircraft; Flight tests; Flow measurement; Speed indicators.


  30. SIMULATOR EVALUATION OF THE LOW-SPEED FLYING QUALITIES OF AN EXPERIMENTAL STOL CONFIGURATION WITH AN EXTERNALLY BLOWN FLAP WING OR AN AUGMENTOR WING
    Authors: B. G. Powers and D. A. Kier
    Report Number: NASA-TN-D-7454
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The low-speed flying qualities of an experimental STOL configuration were evaluated by using a fixed-base six-degree-of-freedom simulation. The configuration had either an externally blown flap (EBF) wing or an augmentor wing (AW). The AW configuration was investigated with two tails, one sized for the AW configuration and a larger one sized for the EBF configuration. The emphasis of the study was on the 70-knot approach task. The stability and control characteristics were compared with existing criteria. Several control systems were investigated for the normal four-engine condition and for the engine-out transient condition. Minimum control and stall speeds were determined for both the three- and four-engine operation.
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    Report Date: October 1973
    No. Pages: 75
    Keywords:      Aerodynamic characteristics; Aerodynamic configurations; Externally blown flaps; Lift augmentation; Short takeoff aircraft.


  31. SUMMARY OF STABILITY AND CONTROL CHARACTERISTICS OF THE XB-70 AIRPLANE
    Authors: C. H. Wolowicz and R. B. Yancey
    Report Number: NASA-TM-X-2933
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The stability and control characteristics of the XB-70 airplane were evaluated for Mach numbers up to 3.0 and altitudes up to 21,300 meters (70,000 feet). The airplane's inherent longitudinal characteristics proved to be generally satisfactory. In the lateral-directional modes, the airplane was characterized by light wheel forces, low static directional stability beyond approximately 2 deg of sideslip, adverse yaw response to aileron inputs throughout the entire Mach number range, and negative effective dihedral with wingtips full down. At subsonic Mach numbers, with the flight augmentation control system off, the light wheel forces and adverse yaw response to aileron inputs caused the pilots to minimize use of the ailerons. At supersonic Mach numbers, with the augmentation system off, the adverse yaw due to aileron and the negative effective dihedral were conducive to pilot-induced oscillations.
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    Subject Category: 08
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    Report Date: October 1973
    No. Pages: 56
    Keywords:      Aerodynamic coefficients; Aircraft performance; Aircraft stability; B-70 aircraft; Flight control.


  32. FLIGHT-DETERMINED STABILITY AND CONTROL CHARACTERISTICS OF THE M2-F3 LIFTING BODY VEHICLE
    Authors: A. G. Sim
    Report Number: NASA-TN-D-7511
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight data were obtained over a Mach number range from 0.4 to 1.55 and an angle-of-attack range from -2 deg to 16 deg. Lateral-directional and longitudinal derivatives, reaction control rocket effectiveness, and longitudinal trim information obtained from flight data and wind-tunnel predictions are compared. The effects of power, configuration change, and speed brake are discussed.
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    Subject Category: 15
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    Report Date: December 1973
    No. Pages: 95
    Keywords:      Aerodynamic configurations; Aircraft performance; Aircraft stability; Flight tests; M-2F3 lifting body.


  33. DEVELOPMENT OF A TURBINE INLET GAS TEMPERATURE MEASUREMENTAND CONTROL SYSTEM USING A FLUIDIC TEMPERATURE SENSOR
    Authors: W. L. Webb and P. J. Reukauf
    Report Number: H-792
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: No abstract available.
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    Subject Category: 07
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    Report Date: November 1973
    No. Pages: 8
    Keywords:      Engine inlets; Fluidic circuits; Gas turbine engines; Temperature measurement; Temperature sensors.
    Notes: American Institute of Aeronautics and Astronautics and Society of Automotive Engineers, Propulsion Conference, 9th, Las Vegas, Nev., Nov. 5-7, 1973. AIAA PAPER 73-1251.


  34. DEVELOPMENT OF AIFTDS-4000, A FLIGHT-QUALIFIED, FLEXIBLE,HIGH-SPEED DATA ACQUISITION SYSTEM
    Authors: R. W. Borek
    Report Number: NASA-TM-X-56018
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NASA flight research center has developed a prototype data acquisition system which integrates an airborne computer with a high-speed pulse code modulation system. The design of the airborne integrated flight test data system (AIFTDS) is the result of experience with airborne pulse code modulation data systems. The AIFTDS-4000 has proved the premise on which it was designed: that the needs and requirements of data acquisition system users can be integrated to produce a highly flexible system that will be more useful than existing systems.
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    Subject Category: 60
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    Report Date: January 1973
    No. Pages: 13
    Keywords:      Airborne/spaceborne computers; Data acquisition; Data processing; Data systems; Pulse code modulation.


  35. SOME STABILITY AND CONTROL ASPECTS OF AIRFRAME/PROPULSIONSYSTEM INTERACTIONS ON THE YF-12 AIRPLANE
    Authors: D. T. Berry and G. B. Gilyard
    Report Number: H-794
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Airframe/propulsion system interactions can strongly affect the stability and control of supersonic cruise aircraft. These interactions generate forces and moments similar in magnitude to those produced by the aerodynamic controls, and can cause significant changes in vehicle damping and static stability. This in turn can lead to large aircraft excursions or high pilot workload, or both. For optimum integration of an airframe and its jet propulsion system, these phenomena may have to be taken into account.
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    Report Date: November 1973
    No. Pages: 7
    Keywords:      Aircraft control; Aircraft stability; Airframes; Fighter aircraft; Jet propulsion.
    Notes: American Society of Mechanical Engineers, Winter Annual Meeting, Detroit, Mich., Nov. 11-15, 1973. ASME PAPER 73-WA/AERO-4.


  36. PRELIMINARY RESULTS OF FLIGHT TESTS OF THE PROPULSION SYSTEM OFTHE YF-12 AIRPLANE AT MACH NUMBERS TO 3.0
    Authors: F. W. Burcham, Jr., J. K. Holzman and P. J. Reukauf
    Report Number: H-795
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests of the propulsion system of a YF-12 airplane were made which included off-schedule inlet operation and deliberately induced unstarts and compressor stalls. The tests showed inlet/engine compatibility to be good through most of the flight envelope. The position of the terminal shock wave could be determined from throat static pressure profiles or from root-mean-square levels of throat static pressure fluctuations. A digital simulation of the control system showed an oscillation of the forward bypass doors to be caused by hysteresis in the bypass door actuator linkages.
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    73
    No. Pages: 12
    Keywords:      Aircraft engines; Engine control; Fighter aircraft; Flight tests; YF-12 aircraft.
    Notes: American Institute of Aeronautics and Astronautics and Society of Automotive Engineers, Propulsion Conference, 9th, Las Vegas, Nev., Nov. 5-7, 1973. AIAA Paper 73-1314.


  37. RESULTS OF A FEASIBILITY STUDY USING THE NEWTON-RAPHSON DIGITAL COMPUTER PROGRAM TO IDENTIFY LIFTING BODY DERIVATIVES FROM FLIGHT DATA
    Authors: A. G. Sim
    Report Number: NASA-TM-X-56017
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A brief study was made to assess the applicability of the Newton-Raphson digital computer program as a routine technique for extracting aerodynamic derivatives from flight tests of lifting body types of vehicles. Lateral-direction flight data from flight tests of the HL-10 lifting body reserch vehicle were utilized. The results in general, show the computer program to be a reliable and expedient means for extracting derivatives for this class of vehicles as a standard procedure. This result was true even when stability augmentation was used. As a result of the study, a credible set of HL-10 lateral-directional derivatives was obtained from flight data. These derivatives are compared with results from wind-tunnel tests
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    Subject Category: 02
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    Report Date: October 1973
    No. Pages: 41
    Keywords:      Aerodynamic stability; Computer programs; HL-10 reentry vehicle; Newton-Raphson method; Research aircraft.


  38. PERFORMANCE OF A BALLUTE DECELERATOR TOWED BEHIND A JET AIRPLANE
    Authors: J. S. Pyle, J. R. Phelps and R. S. Baron
    Report Number: NASA-TM-X-56019
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An F-104B airplane was modified to investigate the drag and stability characteristics of a ballute decelerator in the wake of an asymmetrical airplane. Decelerator deployments were initiated at a Mach number of 1.3 and an altitude of 15,240 meters (50,000 feet) and terminated when the airplane had decelerated to a Mach number of 0.5. The flight tests indicated that the decelerator had a short inflation time with relatively small opening forces. The drag levels attained with the subject decelerator were less than those obtained with other high-speed decelerators behind a symmetrical tow vehicle. The ballute demonstrated good stability characteristics behind the testbed airplane.
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    Report Date: December 1973
    No. Pages: 26
    Keywords:      Aerodynamic drag; Ballutes; Drag devices; F-104 aircraft; Towed bodies.
 
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