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  • 101
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    As Investigation at Mach Numbers of 1.41 and 2.01 of the Aerodynamic Characteristics of an 0.025-scale Model of the MX-1712 (1952)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An investigation of the aerodynamic characteristics of an 0.025-scale model of the MX-1712 configuration has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel. The tests were performed at Mach numbers of 1.41 and 2.01 at a Reynolds number of approximately 2.6 x 10(exp 6) based on the wing mean aerodynamic chord The MX-1712 is a proposed swept-wing, jet-powered supersonic bomber aircraft. The wing is of aspect ratio 3.5, taper ratio 0.2, and thickness ratio 5.5 percent (streamwise) and has 47deg sweep of the quarter-chord line. The longitudinal and lateral force characteristics of the model and various combinations of its components, including several nacelle installations, were investigated. The effects of a modified wing, two horizontal tail positions, and a shortened fuselage were also studied. The results obtained from these investigations are presented in this report. The aerodynamic investigation of this model disclosed no unusual stability characteristics or Mach number effects. The choice of nacelle installations appears to be a major decision, one greatly affecting the performance of the airplane, At M = 1.41 and C(sub L) = 0.1, the buried nacelles increased the drag of the basic model by 9 percent, while the best pod nacelles increased the drag of the basic model by 27 percent.
    Keywords: Aerodynamics
    Type: NACA-RM-SL52J17
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  • 102
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    Heat-transfer and pressure-drop characteristics of internal finned tubes (1951)
    In:  CASI
    Details
    Publication Date: 2019-05-25
    Keywords: unknown
    Type: ASME PAPER-57-A-47
    Format: text
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  • 103
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    Preliminary Investigation of a New Type of Supersonic Inlet (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A supersonic inlet with supersonic deceleration of the flow entirely outside of the inlet is considered. A particular arrangement with fixed geometry having a central body with a circular annular intake is analyzed, and it is shown theoretically that this arrangement gives high pressure recovery for a large range of Mach number and mass flow and therefore is practical for use on supersonic airplanes and missiles. For some Mach numbers the drag coefficient for this type of inlet is larger than the drag coefficient for the type of inlet with supersonic compression entirely inside, but the pressure recovery is larger for all flight conditions. The differences in drag can be eliminated for the design Mach number. Experimental results confirm the results of the theoretical analysis and show that pressure recoveries of 95 percent for Mach numbers of 1.33 and 1.52, 92 percent for a Mach number of 1.72, and 86 percent for a Mach number of 2.10 are possible, with the configurations considered. If the mass flow decreases, the total drag coefficient increases gradually and the pressure recovery does not change appreciably. The results of this work were first presented in a classified document issued in 1946.
    Keywords: Aerodynamics
    Type: NACA-TN-2286
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  • 104
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    Systematic Two-Dimensional Cascade Tests of NACA 65-Series Compressor Blades at Low Speeds (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The performance of NACA 65-series compressor blade section in cascade has been investigated systematically in a low-speed cascade tunnel. Porous test-section side walls and for high-pressure-rise conditions, porous flexible end walls were employed to establish conditions closely simulating two-dimensional flow. Blade sections of design lift coefficients from 0 to 2.7 were tested over the usable angle-of-attack range for various combinations of inlet-flow angle. A sufficient number of combinations were tested to permit interpolation and extrapolation of the data to all conditions within the usual range of application. The results of this investigation indicate a continuous variation of blade-section performance as the major cascade parameters, blade camber, inlet angle, and solidity were varied over the test range. Summary curves of the results have been prepared to enable compressor designers to select the proper blade camber and angle of attack when the compressor velocity diagram and desired solidity have been determined.
    Keywords: Aerodynamics
    Type: NACA-TR-1368 , NACA-RM-L51G31
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  • 105
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    Wind-tunnel investigation of the descent characteristics of bodies of revolution simulating anti-personnel bombs (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation has been conducted in the Langley 20-foot free spinning tunnel to study the relative behavior in descent of a number of homogeneous balsa bodies of revolution simulating anti-personnel bombs with a small cylindrical exploding device suspended approximately 10 feet below the bomb. The bodies of revolution included hemispherical, near-hemispherical, and near-paraboloid shapes. The ordinates of one near-paraboloid shape were specified by the Office of the Chief of Ordnance, U. S. Army. The behavior of the various bodies without the cylinder was also investigated. The results of the investigation indicated that several of the bodies descended vertically with their longitudinal axis, suspension line, and small cylinder in a vertical attitude,. However, the body, the ordinates of which had been specified by the Office of the Chief of Ordnance, U. S. Army, oscillated considerably from a vertical attitude while descending and therefore appeared unsuitable for its intended use. The behavior of this body became satisfactory when its center of gravity was moved well forward from its original position. In general, the results indicated that the descent characteristics of the bodies of revolution become more favorable as their shapes approached that of a hemisphere.
    Keywords: Aerodynamics
    Type: NACA-RM-SL51L13
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  • 106
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    Measurements of downwash and sidewash behind cruciform triangular wings at mach number 1.4 (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Keywords: unknown
    Type: NACA-RM-A51B20
    Format: text
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  • 107
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    Free-flight Investigation of the Longitudinal Stability and Control of a Rocket Propelled Missile Model Having Cruciform, Triangular, Interdigitated Wings and Tails (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Keywords: unknown
    Type: NACA-RM-L51B15
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  • 108
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    A study of the polarization of barium titanate by the transverse electrical field method (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Keywords: unknown
    Type: NASA-CR-92602
    Format: text
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  • 109
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    Determination of Shapes of Boattail Bodies of Revolution for Minimum Wave Drag (1951)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: By use of an approximate equation for the wave drag of slender bodies of revolution in a supersonic flow field, the optimum shapes of certain boattail bodies are determined for minimum wave drag. The properties of three specific families of bodies are determined, the first family consisting of bodies having a given length and base area and a contour passing through a prescribed point between the nose and base, the second family having fixed length, base area, and maximum area, and the third family having given length, volume, and base area. The method presented is easily generalized to determine minimum-wave-drag profile shapes which have contours that must pass through any prescribed number of points. According to linearized theory, the optimum profiles are found to have infinite slope at the nose but zero radius of curvature so that the bodies appear to have pointed noses, a zero slope at the body base, and no variation of wave drag with Mach number. For those bodies having a specified intermediate.diameter (that is, location and magnitude given), the maximum body diameter is shown to be larger, in general, than the specified diameter. It is also shown that, for bodies having a specified maximum diameter, the location of the maximum diameter is not arbitrary but is determined from the ratio of base diameter to maximum diameter.
    Keywords: unknown
    Type: NACA-TN-2550
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  • 110
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    The Effect of Nose Shape on the Drag of Bodies of Revolution at Zero Angle of Attack (1951)
    In:  CASI
    Details
    Publication Date: 2019-08-16
    Description: The subject of this paper is the drag of the nose section of bodies of revolution at zero angle of attack. The magnitude of the nose drag in relation to the total drag is very distinctly a function of the body design and the Mach number. It can range from a very small fraction of the total drag of the order of 10 percent to a very large fraction as high as 80 percent. The natural objective of nose design is to minimize the drag, but this objective is not always the primary one. Sometimes other factors overshadow the desire for minimum drag. The most conspicuous example of this is the proposal of guidance engineers that large-diameter spheres and other very blunt shapes be used at the nose tip. This paper will attempt to discuss both phases of the problem, noses for minimum drag and noses with very blunt tips. The state of the theory will also be reviewed and recent theoretical developments described, since the theory still remains a very valuable tool for assaying the effects of compromises in design and departure from shapes for which experimental data are available.
    Keywords: Aerodynamics
    Type: Aerodynamic Characteristics of Bodies at Supersonic Speeds: A Collection of Three Papers; 1-12; NACA-RM-A51J25|NACA Conference on Aerodynamic Design Problems of Supersonic Guided Missiles; Oct 02, 1951 - Oct 03, 1951; Moffett Field, CA; United States
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  • 111
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    An Investigation of the Aerodynamic Characteristics in Pitch of an XB-52 Airplane Model in a High-Speed Wind Tunnel (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A wind-tunnel investigation of a 0.049-scale model of the Boeing XB-52 airplane was made at Mach numbers from 0.30 to 0.925 and at corresponding Reynolds numbers from about 2.3 x 10(exp 6) to 4.3 x 10(exp 6). The results of the investigation indicate satisfactory static longitudinal stability throughout the test Mach-number range and some loss in tail effectiveness beginning at about 0.80 Mach number. A comparison of the results of these tests with those of the same model in the Boeing Airplane Company's wind tunnel showed close agreement of lift- and drag-divergence Mach numbers. Slight differences were observed in tail effectiveness and the position of the stick-fixed neutral point.
    Keywords: Aerodynamics
    Type: NACA-RM-SA51C16
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  • 112
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    Supplementary Free-Spinning-Tunnel Investigation of a 1/30-Scale Model of the Grumman XF10F-1 Airplane in the Swept-Wing Configuration with Slats Extended (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A supplementary investigation has been conducted in the Langley 20-foot free-spinning tunnel of a 1/30 -scale model of the Grumman XFlOF-1 airplane to determine what effect full-span slats would have on the spin-recovery characteristics of the swept-wing version of the XFlOF-1 airplane, which had previously been indicated as possessing undesirable spin-recovery characteristics without slats. The effects of extended nose-wheel doors and of fairing the air-duct inlets were also determined. The results indicated that, with slats fully extended, satisfactory recovery could be obtained by rudder reversal provided it was accompanied by movement of the trimmer ailerons to full with the spin (only up-going spoiler operative), Extension of the nose-wheel doors or fairing of the air-duct inlets did not improve the recovery characteristics.
    Keywords: Aerodynamics
    Type: NACA-RM-SL51G19
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  • 113
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    Damping in a Roll of a Missile Configuration with a Modified Triangular Wing and a Cruciform Tail at a Mach Number of 1.52 (1951)
    In:  CASI
    Details
    Publication Date: 2019-08-14
    Description: The damping-in-Toll stability derivatives of a missile configuration and its components were determined both experimentally and theoretically. The tests were conducted at a Mach number of 1.52 and at a Reynolds number, based on the mean aerodynamic chord of the wing, of 0.82 x 10(exp 6). The experimental damping derivative of the wing-body combination was 67 percent of the theoretical value. The difference is believed to have resulted mainly from the fact that the theory is not strictly applicable when the Mach number normal to the leading edge is almost unity, which was the case in the present investigation. For the tail-body combination the damping derivative was 86 percent of the theoretical value. In this case, the difference is believed to have been caused partially by mutual interference between the tail surfaces and partially by the low Reynolds number of the flow over the tail. It was found that the damping of the complete configuration was not equal to the sum of the damping derivatives of the components because of the effect of the wing downwash on the damping of the tail.
    Keywords: Aerodynamics
    Type: NACA-RM-A51A03
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  • 114
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    Lift, Drag, and Pitching Moment of Low-Aspect-Ratio Wings at Subsonic and Supersonic Speeds: Triangular Wing of Aspect Ratio 2 with NACA 0005-63 Thickness Distribution, Cambered and Twisted for a Trapezoidal Span Load Distribution (1951)
    In:  CASI
    Details
    Publication Date: 2019-08-17
    Description: A wing-body combination having a plane triangular wing of aspect ratio 2 with NACA 0005-63 thickness distribution in streamwise planes, and twisted and cambered for a trapezoidal span load distribution has been investigated at both subsonic and supersonic Mach numbers. The lift, drag, and pitching moment of the model are presented for Mach numbers from 0.60 to 0.90 and 1.30 to 1.70 at a Reynolds number of 3.0 million. The variations of the characteristics with Reynolds number are also shown for several Mach numbers.
    Keywords: Aerodynamics
    Type: NACA-RM-A50K27a
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  • 115
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    Aerodynamic Characteristics of a Flying-Boat Hull Having a Length-Beam Ratio of 15, TED No. NACA 2206 (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the aerodynamic characteristics of a flying-boat hull of a length-beam ratio of 15 in the presence of a wing. The investigation was an extension of previous tests made on hulls of length-beam ratios of 6, 9, and 12; these hulls were designed to have approximately the same hydrodynamic performance with respect to spray and resistance characteristics. Comparison with the previous investigation at lower length-beam ratios indicated a reduction in minimum drag coefficients of 0.0006 (10 peroent)with fixed transition when the length-beam ratio was extended from 12 to 15. As with the hulls of lower length-beam ratio, the drag reduction with a length-beam ratio of 15 occurred throughout the range of angle of attack tested and the angle of attack for minimum drag was in the range from 2deg to 3deg. Increasing the length-beam ratio from 12 to 15 reduced the hull longitudinal instability by an mount corresponding to an aerodynamic-center shift of about 1/2 percent of the mean aerodynamic chord of the hypothetical flying boat. At an angle of attack of 2deg, the value of the variation of yawing-moment coefficient with angle of yaw for a length-beam ratio of 15 was 0.00144, which was 0.00007 larger than the value for a length-beam ratio of 12.
    Keywords: Aerodynamics
    Type: NACA-RM-L6J24
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  • 116
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    Large-Scale Flight Measurements of Zero-Lift Drag at Mach Numbers from 0.87 to 1.39 of 1/10-Scale Models of the Northrop MX-775A Missile (1951)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: A flight investigation was made at high subsonic, transonic, and supersonic speeds and at high Reynolds numbers to determine the zero-lift drag of a 1/10-scale model of the Northrop MX-775A missile and a scale model of the missile fuselage. The model of the complete configuration has a 45deg swept wing of aspect ratio 5.5 and a 33deg swept vertical fin. The body model was stabilized by three 45deg swept fins. The-drag-rise Mach number for the model of the complete configuration was approximately 0.96. The drag coefficient based on total wing area was 0.0330 at Mach number 1.39. The drag coefficient of the body model less fin drag was approximately 55 percent that of the complete model at the same Mach number. Addition of the wing to the fuselage apparently resulted in a favorable drag interference near Mach number 1.0.
    Keywords: Aerodynamics
    Type: NACA-RM-SL51K07
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  • 117
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    Base Pressure on Wings and Bodies with Turbulent Boundary Layers (1951)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: At present there is no satisfactory theory for calculating the pressure which acts at the blunt base of an object traveling at supersonic velocity. In fact, the essential mechanism determining the base pressure is only imperfectly understood. As a result, the existing knowledge of base pressure is based almost entirely on experiments. The main object of this paper is to summarize the principal results of the many wind tunnel and free flight measurements of base pressure on both bodies of revolution and blunt trailing edge airfoils. A relatively simple method of estimating base pressure is presented, and an indication is given as to how the characteristics of base pressure play an essential role in determining the shape of an aerodynamically efficient object for supersonic flight.
    Keywords: Aerodynamics
    Type: Aerodynamic Characteristics of Bodies at Supersonic Speeds: A Collection of Three Papers; 13-30; NACA-RM-A51J25|NACA Conference on Aerodynamic Design Problems of Supersonic Guided Missiles; Oct 02, 1951 - Oct 03, 1951; Moffett Field, CA; United States
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  • 118
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    Aerodynamic Characteristics of Bodies at Supersonic Speeds: A Collection of Three Papers (1951)
    In:  CASI
    Details
    Publication Date: 2019-08-15
    Description: The three papers collected here are: 'The Effect of Nose Shape on the Drag of Bodies of Revolution at Zero Angle of Attack.', 'Base Pressure on Wings and Bodies with Turbulent Boundary Layers', and 'Flow over Inclined Bodies'. The subject of the first paper is the drag of the nose section of bodies of revolution at zero angle of attack. The main object of the second paper is to summarize the prinicpal results of the many wind tunnel and free flight measurements of base pressure on both bodies of revolution and blunt trailing edge airfoils.
    Keywords: Aerodynamics
    Type: NACA-RM-A51J25
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  • 119
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    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 turn at a mach number of 0.86 (1950)
    In:  CASI
    Details
    Publication Date: 2019-05-25
    Keywords: unknown
    Type: NACA-RM-L50J10
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  • 120
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    An Analysis of the Applicability of the Hypersonic Similarity Law to the Study of Flow About Bodies of Revolution at Zero Angle of Attack (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The hypersonic similarity law as derived by Tsien has been investigated by comparing the pressure distributions along bodies of revolution at zero angle of attack. In making these comparisons, particular attention was given to determining the limits of Mach number and fineness ratio for which the similarity law applies. For the purpose of this investigation, pressure distributions determined by the method of characteristics for ogive cylinders for values of Mach numbers and fineness ratios varying from 1.5 to 12 were compared. Pressures on various cones and on cone cylinders were also compared in this study. The pressure distributions presented demonstrate that the hypersonic similarity law is applicable over a wider range of values of Mach numbers and fineness ratios than might be expected from the assumptions made in the derivation. This is significant since within the range of applicability of the law a single pressure distribution exists for all similarly shaped bodies for which the ratio of free-stream Mach number to fineness ratio is constant. Charts are presented for rapid determination of pressure distributions over ogive cylinders for any combination of Mach number and fineness ratio within defined limits.
    Keywords: Aerodynamics
    Type: NACA-TN-2250
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  • 121
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    An Approximate Method of Calculating Pressures in the Tip Region of a Rectangular Wing of Circular-Arc Section at Supersonic Speeds (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: No abstract available
    Keywords: Aerodynamics
    Type: NACA-TN-2211
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  • 122
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    Wind-Tunnel Investigation of a 0.6-Scale Model of Hughes MX-904 Tail Surface at Supersonic Speeds. Tail Attached to a Segment of the Foreshortened Body (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Keywords: unknown
    Type: NACA-RM-SL50E10
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  • 123
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    Reaction kinetics of rocket propellant gases - 1. rate of decomposition of nitric oxide at elevated temperatures (1950)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Keywords: unknown
    Type: PR-9-46 , PB-169143
    Format: text
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  • 124
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    Model ditching investigations of three airplanes equipped with hydro-skis (1950)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Keywords: unknown
    Type: NASA-TM-X-61112 , NACA-RM-L9K23
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  • 125
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    The accuracy of present wave forecasting methods with reference to problems in beach erosion on the New Jersey and Long Island coasts (1950)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Keywords: unknown
    Type: NASA-CR-101808
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  • 126
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    Large-Scale Flight Measurements of Zero-Lift Drag at Mach Numbers from 0.90 to 1.95 of 1/14-Scale Model of the Northrop MX-775B Pilotless Aircraft with Small Body (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A flight test was made a t high subsonic, transonic, and supersonic speeds and at high Reynolds numbers to determine the zero-lift drag of a 1/14-scale model of the Northrop MX-775B pilotless aircraft with small small body. The triangular wing of the model had 67.5 deg leading-edge sweep and 15 deg. trailing-edge sweep, The wing airfoil sections were modified NACA 0004 sections. The drag coefficient based on total wing area was 0.0107 at Mach number 1.60. At transonic speeds the maximum drag coefficient was 0.0125. The force-break Mach number was 0,98.
    Keywords: Aerodynamics
    Type: NACA-RM-SL50H18
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  • 127
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    Tests of a Hermes A-2 Missile Body at Mach Number 4.04 (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: Force tests on a proposed body shape of the Hermes A-2 missile with and without longitudinal spoilers were made at Mach number 4.04. Values of normal force coefficient, pitching-moment coefficient, and center-of-pressure position were obtained.
    Keywords: Aerodynamics
    Type: NACA-RM-SL50H23A
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  • 128
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    Free-Spinning-Tunnel Investigation of a 1/24-Scale Model of the Grumman AF-2S, -2W Airplane (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation of the spin and recovery characteristics of a 1/24-scale model of the Grumman AF-2S, -2W airplane was conducted in the Langley 20-foot free-spinning tunnel. The effects of controls on the erect and inverted spin and recovery characteristics for a range of possible loadings of the.airplane were determined. The effect of a revised-tail installation (small dual fins added to the stabilizer of the original tail and the vertical-tail height of the original tail increased) and the effect of various ventral-fin and antispin-fillet installations were determined. The investigation also included spin-recovery parachute tests.
    Keywords: Aerodynamics
    Type: NACA-RM-SL51B20
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  • 129
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    Wind-Tunnel Investigation of a 0.6-Scale Model of Hughes MX-904 Tail Surface at Supersonic Speeds Several Combinations of the Tail with Each of Two Foreshortened Body Segments (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: An investigation has been made in the Langley 9- by 12-inch super-sonic blowdown tunnel at Mach numbers of 1.62 and 1.96 of a partial-span body with one tail surface, designed for use on the Hughes Falcon (MX-904) missile. The present paper extends the work reported in NACA-RM-SL50E10. Force and moment data including elevator hinge moment were obtained for the conditions of the tail in the presence of a small segment of the fore-shortened body, in the presence of a semi-span body and attached to a semi-span body, and for the condition of the foreshortened semi-span body alone.
    Keywords: Aerodynamics
    Type: NACA-RM-SL50G13
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  • 130
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    Free-Spinning-Tunnel Investigation of a 1/30-Scale Model of the Grumman XF1OF-1 Airplane (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: An investigation has been conducted in the Langley 20-foot free-spinning tunnel on a 1/30 - scale model of the Grumman XFlOF-1 airplane to determine its spin and recovery characteristics. The investigation included erect and inverted spins for both the straight-wing and swept-wing configurations. Tests to determine the optimum size spin-recovery parachutes and the rudder forces required for recovery were also made. The results indicated that in the straight-wing configuration, satisfactory recoveries of the airplane will be obtained from erect and inverted spins by rudder reversal alone. In the swept-wing configuration recoveries will be unsatisfactory from erect spins. Unsweeping the wings during the spin and reversal of the rudder, however, will lead to eventual recovery. The test results also indicated that, if existing small ailerons are made deflectable through large angles, satisfactory recoveries will be obtained from erect spins in the swept-wing configuration by simultaneous movement of the rudder to against the spin and movement of the ailerons to with the spin. Normal-size ailerons deflected through a normal range would also be effective. Satisfactory recoveries by rudder reversal will be obtained from inverted spins in the swept-wing configuration. In the straight-wing configuration a 14.2-foot tail parachute or a 5.0-foot wing-tip parachute opened on the outer wing tip will effect satisfactory recovery of the airplane by parachute action alone; a 30.0-foot tail parachute or a 10.0-foot wing-tip parachute will be required for the swept-wing configuration. The forces required to fully reverse the rudder should be within the capabilities of the pilot.
    Keywords: Aerodynamics
    Type: NACA-RM-SL50L14
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  • 131
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    Dynamic Response of Control Servo System Installed in NAES-Equipped SB2C-5 Airplane (BuAer No. 83135) (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: Dynamic--response measurements for various conditions of displacement and rate signal input, sensitivity setting, and simulated hinge moment were made of the three control-surface servo systems of an NAES-equipped remote-controlled airplane while on the ground. The basic components of the servo systems are those of the General Electric Company type G-1 autopilot using electrical signal. sources, solenoid-operated valves, and hydraulic pistons. The test procedures and difficulties are discussed, Both frequency and transient-response data, are presented and comparisons are made. The constants describing the servo system, the undamped natural frequency, and the damping ratio, are determined by several methods. The response of the system with the addition of airframe rate signal is calculated. The transfer function of the elevator surface, linkage, and cable system is obtained. The agreement between various methods of measurement and calculation is considered very good. The data are complete enough and in such form that they may be used directly with the frequency-response data of an airplane to predict the stability of the autopilot-airplane combination.
    Keywords: Aerodynamics
    Type: NACA-RM-SA50J05
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  • 132
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    Preliminary Performance Data on Westinghouse Electronic Power Regulator Operating on J34-WE-32 Turbojet Engine in Altitude Wind Tunnel (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: The behavior of the Westinghouse electronic power regulator operating on a J34-WE-32 turbojet engine was investigated in the NACA Lewis altitude wind tunnel at the request of the Bureau of Aeronautics, Department of the Navy. The object of the program was to determine the, steady-state stability and transient characteristics of the engine under control at various altitudes and ram pressure ratios, without afterburning. Recordings of the response of the following parameters to step changes in power lever position throughout the available operating range of the engine were obtained; ram pressure ratio, compressor-discharge pressure, exhaust-nozzle area, engine speed, turbine-outlet temperature, fuel-valve position, jet thrust, air flow, turbine-discharge pressure, fuel flow, throttle position, and boost-pump pressure. Representative preliminary data showing the actual time response of these variables are presented. These data are presented in the form of reproductions of oscillographic traces.
    Keywords: Aerodynamics
    Type: NACA-RM-SE50J11
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  • 133
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    An Investigation of the Longitudinal Characteristics of the MX-656 Configuration Using Rocket-Propelled Models Preliminary Results at Mach Numbers from 0.65 to 1.25 (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-12
    Description: A rocket-propelled model of the Mx-656 configuration has been flown through the Mach number range from 0.65 to 1.25. An analysis of the response of the model to rapid deflections of the horizontal tail gave information on the lift, drag, longitudinal stability and control, and longitudinal-trim change. The lift-coefficient range covered by the test was from -0.2 to 0,3 throughout most of the Mach number range, The model was statically and dynamically stable throughout the lift-coefficient and Mach number range of the test. At subsonic speeds the aerodynamic center moved f o m r d with increasing lift coefficient. The most forward position of the aerodynamic center was about 12,5 percent of the mean aerodynamic chord at a small positive lift coefficient and at a Mach number of about 0.84. A t supersonic speeds the aerodynamic center was well aft, varying from 33 to 39 percent of the mean aerodynamic chord at Mach numbers of 1.0 and 1.25, respectively. Transonic-trim change, as measured by the change in trim lift coefficient with Mach number at a constant t a i l setting, was of small magnitude (about 0.1 lift coefficient for zero tail setting). The zero-lift/drag coefficient increased about 0.042 in the region between a Mach number of 0.9 and 1.1
    Keywords: Aerodynamics
    Type: NACA-RM-SL50J03
    Format: application/pdf
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  • 134
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    Investigation of the Model ME 210 in the Spin Wind Tunnel of the DVL (1950)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: After conclusion of the spin investigation of the model Me 210 with elongated fuselage and central vertical tail surfaces (model condition III; reference 3), tests were performed on the same model with a vee tail (model condition IV). Here the entire tail surfaces consist of only one surface with pronounced dihedral. Since the blanketing of the vertical tail surfaces by the horizontal tail surfaces, which may occur in case of standard tail surfaces, does not occur here, one could expect for this type of tail surface favorable spin characteristics, particularly with respect to rudder effectiveness for spin recovery. However, the test results did not confirm these expectations. The steady spin was shown to be very irregular; regarding rudder effectiveness the vee tail surfaces proved to be inferior even to standard tail surfaces, thus they represent the most unfavorable of the four fuselage and tail-surface combinations investigated so far.
    Keywords: Aerodynamics
    Type: NACA-TM-1222 , Zentrale fuer Wissenschaftliches Berichtswesen der Luftfahrtforschung des Generalluftzeugmeisters (ZWB) Untersuchungen und Mitteilungen; Rept-1288
    Format: application/pdf
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