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  • 1
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    Preliminary Wind-Tunnel Investigation of the Performance of Republic F-105 Wing-Root Inlet Configurations at Various Angles of Attack and a Mach Number of 2.01 (1957)
    In:  CASI
    Details
    Publication Date: 2019-05-31
    Description: A 1/13-scale model of the forebody of the Republic F-105 with twin-duct wing-root inlets was tested in the Langley 4- by 4-foot supersonic pressure tunnel through a range of angle of attack from -4 deg to 15 deg at a Mach number of 2.01 and a Reynolds number of approximately 3.4 x 10(exp 6) per foot. The tests were made with four configurations which incorporated varying amounts of sweep and stagger of the inlet leading edges, modifications to the areas of the boundary-layer diverter floor plate, and modifications to the area of the boundary-layer diverter bleed slots. The highest overall pressure recovery at an angle of attack of 0 deg (average total-pressure recovery, 0.84 mass-flow ratio, 0.98) was achieved with configuration having an inlet leading-edge sweep angle of 58 deg with no leading-edge stagger. Stagger was found to improve the angle-of- attack performance, but at a sacrifice in inlet efficiency for an angle of attack of 0 deg. The boundary-layer diverter floor height, of the order of one boundary-layer thickness, was satisfactory for bypassing the fuselage boundary layer. The boundary-layer diverter-plate bleed slots were effective in increasing the total-pressure recovery of the inlet. The total-pressure-recovery contour plots, taken at the compressor-face station, indicate the existence of high-velocity "cores" throughout the inlet operating range.
    Keywords: Aerodynamics
    Type: NACA-RM-SL56L12
    Format: application/pdf
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  • 2
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    Design Guide for Pitch-Up Evaluation and Investigation at High Subsonic Speeds of Possible Limitations Due to Wing-Aspect-Ratio Variations (1959)
    In:  CASI
    Details
    Publication Date: 2019-05-11
    Description: A design guide is suggested as a basis for indicating combinations of airplane design variables for which the possibilities of pitch-up are minimized for tail-behind-wing and tailless airplane configurations. The guide specifies wing plan forms that would be expected to show increased tail-off stability with increasing lift and plan forms that show decreased tail-off stability with increasing lift. Boundaries indicating tail-behind-wing positions that should be considered along with given tail-off characteristics also are suggested. An investigation of one possible limitation of the guide with respect to the effects of wing-aspect-ratio variations on the contribution to stability of a high tail has been made in the Langley high-speed 7- by 10-foot tunnel through a Mach number range from 0.60 to 0.92. The measured pitching-moment characteristics were found to be consistent with those of the design guide through the lift range for aspect ratios from 3.0 to 2.0. However, a configuration with an aspect ratio of 1.55 failed t o provide the predicted pitch-up warning characterized by sharply increasing stability at the high lifts following the initial stall before pitching up. Thus, it appears that the design guide presented herein might not be applicable when the wing aspect ratios lower than about 2.0.
    Keywords: Aerodynamics
    Type: NASA-TM-X-26
    Format: application/pdf
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  • 3
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    An Investigation of Some Aspects of the Sonic Boom by Means of Wind-Tunnel Measurements of Pressures About Several Bodies at a Mach Number of 2.01 (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An investigation of some aspects of the sonic boom has been made with the aid of wind-tunnel measurements of the pressure distributions about bodies of various shapes. The tests were made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.01 and at a Reynolds number per foot of 2.5 x 10(exp 6). Measurements of the pressure field were made at orifices in the surface of a boundary-layer bypass plate. The models which represented both fuselage and wing types of thickness distributions were small enough to allow measurements as far away as 8 body lengths or 64 chords. The results are compared with estimates made using existing theory. To the first order, the boom-producing pressure rise across the bow shock is dependent on the longitudinal development of body area and not on local details. Nonaxisymmetrical shapes may be replaced by equivalent bodies of revolution to obtain satisfactory theoretical estimates of the far-field pressures.
    Keywords: Aerodynamics
    Type: NASA-TN-D-161
    Format: application/pdf
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  • 4
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    Ground Measurements of the Shock-Wave Noise from Airplanes in Level Flight at Mach Numbers to 1.4 and at Altitudes to 45000 Feet (1959)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Time histories of noise pressures near ground level were measured during flight tests of fighter-type airplanes over fairly flat, partly wooded terrain in the e Mach number range between 1.13 and 1.4 and at altitudes from 25,000 to 45,000 feet. Atmospheric soundings and radar tracking studies were made for correlation with the measured noise data. The measured and calculated values of the pressure rise across the shock wave were generally in good agreement. There is a tendency for the theory to overestimate the pressure at locations remote from the track and to underestimate the pressures for conditions of high tailwind at altitude. The measured values of ground-reflection factor averaged about 1.8 f or the surface tested as compared to a theoretical value of 2.0. P o booms were measured in all cases. The observers also generally reported two booms; although, in some cases, only one boom was reported. The shock-wave noise associated with some of the flight tests was judged to be objectionable by ground observers, and in one case the cracking of a plate-glass store window was correlated in time with the passage of the airplane at an altitude of 25,000 feet.
    Keywords: Aerodynamics
    Type: NASA-TN-D-48
    Format: application/pdf
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  • 5
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    Exploratory Wind-Tunnel Investigation to Determine the Lift Effects of Blowing over Flaps from Nacelles Mounted Above the Wing (1958)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An exploratory wind-tunnel investigation has been made to determine the lift effects of blowing from nacelles over the upper surface of flaps on a model having a delta wing of aspect ratio 3. Several flap conditions were examined. High-pressure air was blown from an external-pipe arrangement supported above the wing to simulate jet-engine exhaust. The jet momentum- coefficient range was from 0 to 3.0 and the model angle of attack was 0 deg. The results of this limited investigation show that values of jet circulation lift coefficient larger than the Jet reaction were produced with blowing over flaps from nacelles mounted above the wing. 'I!heuse of double slotted flaps with the gap unsealed between the flaps and wing had a large detrimental effect on the lift capabilities. With these gaps sealed, larger lift coefficients were obtained when fantails were added to the nacelles. The longitudinal trim problems created by large diving moments were similar to those encountered with other jet-augmented-flap systems
    Keywords: Aerodynamics
    Type: NACA-TN-4298
    Format: application/pdf
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  • 6
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    Theoretical Calculations of Supersonic Wave Drag at Zero Lift for a Particular Store Arrangement (1958)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An analysis, based on the linearized thin-airfoil theory for supersonic speeds, of the wave drag at zero lift has been carried out for a simple two-body arrangement consisting of two wedgelike surfaces, each with a rhombic lateral cross section and emanating from a common apex. Such an arrangement could be used as two stores, either embedded within or mounted below a wing, or as auxiliary bodies wherein the upper halves could be used as stores and the lower halves for bomb or missile purposes. The complete range of supersonic Mach numbers has been considered and it was found that by orienting the axes of the bodies relative to each other a given volume may be redistributed in a manner which enables the wave drag to be reduced within the lower supersonic speed range (where the leading edge is substantially subsonic). At the higher Mach numbers, the wave drag is always increased. If, in addition to a constant volume, a given maximum thickness-chord ratio is imposed, then canting the two surfaces results in higher wave drag at all Mach numbers. For purposes of comparison, analogous drag calculations for the case of two parallel winglike bodies with the same cross-sectional shapes as the canted configuration have been included. Consideration is also given to the favorable (dragwise) interference pressures acting on the blunt bases of both arrangements.
    Keywords: Aerodynamics
    Type: NACA-TN-4120
    Format: application/pdf
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  • 7
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    A study of the motion and aerodynamic heating of missiles entering the earth's atmosphere at high supersonic speeds (1957)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A simplified analysis of the velocity and deceleration history of missiles entering the earth's atmosphere at high supersonic speeds is presented. The results of this motion analysis are employed to indicate means available to the designer for minimizing aerodynamic heating. The heating problem considered involves not only the total heat transferred to a missile by convection, but also the maximum average and local time rates of convective heat transfer.
    Keywords: Aerodynamics
    Type: NACA-TN-4047
    Format: application/pdf
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  • 8
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    The Compressible Laminar Boundary Layer with Fluid Injection (1955)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A solution of the equations of the compressible laminar boundary layer including the effects of transpiration cooling is presented. The analysis applies to the flow over an isothermal porous plate with a velocity of fluid injection proportional to the reciprocal of the square root of the distance from the leading edge. The effect of several flow parameters on coolant-flow rates is discussed with the aid of representative examples. A stability analysis indicates that, although transpiration cooling requires a lower surface temperature for stable flow than does internal wall cooling, this lower temperature can be obtained with a smaller expenditure of coolant.
    Keywords: Aerodynamics
    Type: NACA-TN-3404
    Format: application/pdf
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  • 9
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    Aerodynamic Design of Axial-flow Compressors (1956)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Available experimental two-dimensional-cascade data for conventional compressor blade sections are correlated. The two-dimensional cascade and some of the principal aerodynamic factors involved in its operation are first briefly described. Then the data are analyzed by examining the variation of cascade performance at a reference incidence angle in the region of minimum loss. Variations of reference incidence angle, total-pressure loss, and deviation angle with cascade geometry, inlet Mach number, and Reynolds number are investigated. From the analysis and the correlations of the available data, rules and relations are evolved for the prediction of the magnitude of the reference total-pressure loss and the reference deviation and incidence angles for conventional blade profiles. These relations are developed in simplified forms readily applicable to compressor design procedures.
    Keywords: Aerodynamics
    Type: NACA-RM-E56B03a
    Format: application/pdf
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  • 10
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    Some Effects of Roll Rate on the Longitudinal Stability Characteristics of a Cruciform Missile Configuration as Determined from Flight Test for a Mach Number Range of 1.1. to 1.8 (1956)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: A model of a cruciform missile configuration having a low-aspect-ratio wing equipped with flap-type controls was flight tested in order to determine stability and control characteristics while rolling at about 5 radians per second. Comparison is made with results from a similar model which rolled at a much lower rate. Results showed that, if the ratio of roll rate to natural circular frequency in pitch is not greater than about 0.3, the motion following a step disturbance in pitch essentially remains in a plane in space. The slope of normal- force coefficient against angle of attack C(sub N(sub alpha)) was the same as for the slowly rolling model at 0 degrees control deflection but C(sub N(sub alpha)) was much higher for the faster rolling model at about 5 degrees control deflection. The slope of pitching-moment coefficient against angle of attack C(sub m(sub alpha)) as determined from the model period of oscillation was the same for both models at 0 degrees control deflection but was lower for the faster rolling model at about 5 degrees control deflection. Damping data for the faster rolling model showed considerably more scatter than for the slowly rolling model.
    Keywords: Aerodynamics
    Type: NACA-RM-L55L16
    Format: application/pdf
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