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At least 199 records · Page 11

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NASA NTRS1978

Low-speed aerodynamic characteristics of a 16-percent-thick variable-geometry airfoil designed for general aviation applications

Tests were conducted in the Langley low-turbulence pressure tunnel to determine the aerodynamic characteristics of climb, cruise, and landing configurations. These tests were conducted over a Mach number range from 0.10 to 0.35, a chord Reynolds number range from 2.0 x 1 million to 20.0 x 1 million, and an angle-of-attack range from -8 deg to 20 deg. Results show that the maximum section lift coefficients increased in the Reynolds number range from 2.0 x 1 million to 9.0 x 1 million and reached values of approximately 2.1, 1.8, and 1.5 for the landing, climb, and cruise configurations, respectively. Stall characteristics, although of the trailing-edge type, were abrupt. The section lift-drag ratio of the climb configuration with fixed transition near the leading edge was about 78 at a lift coefficient of 0.9, a Mach number of 0.15, and a Reynolds number of 4.0 x 1 million. Design lift coefficients of 0.9 and 0.4 for the climb and cruise configurations were obtained at the same angle of attack, about 6 deg, as intended. Good agreement was obtained between experimental results and the predictions of a viscous, attached-flow theoretical method.

Barnwell, R. W.↗

NASA NTRS1994

Interpretation of waverider performance data using computational fluid dynamics

A computational study was conducted to better understand experimental results obtained from wind-tunnel tests of a Mach 4 conical-flow-derived waverider and a comparative reference configuration, which showed that the aerodynamic performance of the reference configuration was slightly better than that of the waverider. The computational results showed that the predicted surface pressure values and the integrated lift and drag coefficients were much lower for the reference model because the reference model bottom is an expansion surface. However, the lift-drag ratios for the reference model were higher due to a relatively low drag for a comparable amount of lift. The results also showed that the reference model exhibited the same shock attachment characteristics as the conical-flow-derived waverider, and is therefore also a waverider. The shock attachment characteristic gives the waverider a performance advantage over conventional hypersonic vehicles, and the results suggest that altering the bottom surface does not cause significant performance degradation. Flowfield solutions also show that the conical-flow waverider model has better propulsion/airframe integration characteristics than the reference configuration. The results also suggest that generating flowfields other than conical ones may be used to design waveriders with improved aerodynamic performance.

COMPUTATIONAL FLUID DYNAMICS↗

NASA NTRS1986

Over-the-wing propeller

This invention is an aircraft with a system for increasing the lift drag ratio over a broad range of operating conditions. The system positions the engines and nacelles over the wing in such a position that gains in propeller efficiency is achieved simultaneously with increases in wing lift and a reduction in wing drag. Adverse structural and torsional effects on the wings are avoided by fuselage mounted pylons which attach to the upper portion of the fuselage aft of the wings. Similarly, pylon-wing interference is eliminated by moving the pylons to the fuselage. Further gains are achieved by locating the pylon surface area aft of the aircraft center of gravity, thereby augmenting both directional and longitudinal stability. This augmentation has the further effect of reducing the size, weight and drag of empennage components. The combination of design changes results in improved cruise performance and increased climb performance while reducing fuel consumption and drag and weight penalties.

Johnson, Joseph L., Jr.↗

NASA NTRS1973

Measured pressure distributions of large-angle cones in hypersonic flows of tetrafluoromethane, air, and helium

An experimental study of surface pressure distributions on a family of blunt and sharp large angle cones was made in hypersonic flows of helium, air, and tetrafluoromethane. The effective isentropic exponents of these flows were 1.67, 1.40, and 1.12. Thus, the effect of large shock density ratios such as might be encountered during planetary entry because of real-gas effects could be studied by comparing results in tetrafluoromethane with those in air and helium. It was found that shock density ratio had a large effect on both shock shape and pressure distribution. The differences in pressure distribution indicate that for atmospheric flight at high speed where real-gas effects produce large shock density ratios, large-angle cone vehicles can be expected to experience different trim angles of attack, drag coefficient, and lift-drag ratios than those for ground tests in air wind tunnels.

Jones, R. A.↗

NASA NTRS1979

Transonic aerodynamic characteristics of a supersonic cruise aircraft research model with the engines suspended above the wing

The influence of upper-surface nacelle exhaust flow on the aerodynamic characteristics of a supersonic cruise aircraft research configuration was investigated in a 16 foot transonic tunnel over a range of Mach numbers from 0.60 to 1.20. The arrow-wing transport configuration with engines suspended over the wing was tested at angles of attack from -4 deg to 6 deg and jet total pressure ratios from 1 to approximately 13. Wing-tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration. Various nacelle locations (chordwise, spanwise, and vertical) were tested over the ranges of Mach numbers, angles of attack, and jet total-pressure ratios. The results show that reflecting the wing-tip leading edge flap from 0 deg to -10 deg increased the maximum lift-drag ratio by 1.0 at subsonic speeds. Jet exhaust interference effects were negligible.

Mercer, C. E.↗

NASA NTRS1955

Optimum flight paths of turbojet aircraft

The climb of turbojet aircraft is analyzed and discussed including the accelerations. Three particular flight performances are examined: minimum time of climb, climb with minimum fuel consumption, and steepest climb. The theoretical results obtained from a previous study are put in a form that is suitable for application on the following simplifying assumptions: the Mach number is considered an independent variable instead of the velocity; the variations of the airplane mass due to fuel consumption are disregarded; the airplane polar is assumed to be parabolic; the path curvatures and the squares of the path angles are disregarded in the projection of the equation of motion on the normal to the path; lastly, an ideal turbojet with performance independent of the velocity is involved. The optimum Mach number for each flight condition is obtained from the solution of a sixth order equation in which the coefficients are functions of two fundamental parameters: the ratio of minimum drag in level flight to the thrust and the Mach number which represents the flight at constant altitude and maximum lift-drag ratio.

CLIMBING SPEEDS - EFFECT OF ALTITUDE↗

NASA NTRS1959

An Experimental Investigation of a Triangular Wing of Aspect Ratio 2 and a Body Warped to be Trimmed at M = 2.24

A cambered and twisted triangular wing of aspect ratio 2 in combination with a cambered body was investigated experimentally to determine the effectiveness of the camber in reducing the drag due to lift at trim at supersonic speeds. Four arrangements were tested comprising all combinations of a symmetrical and a cambered wing with a symmetrical and a cambered body. The camber shape investigated was derived by linearized lifting surface theory for triangular wings with sonic leading edges and satisfied the requirement that the wing be trimmed at the design Mach number and lift coefficient. The experimental results for the cambered wing and cambered body showed that the drag coefficient at trim was always greater, at the same lift coefficient, than that for the untrimmed symmetrical wing and body. The trim lift coefficient was positive and decreased with increasing Mach number. At the design Mach number of 2.24, the trim lift coefficient was somewhat lower and the drag coefficient was higher than values predicted by linearized lifting surface theory for the wing alone. A comparison of the trim lift-drag ratio of the cambered wing and cambered body with values obtained by trimming the symmetrical wing and symmetrical body either with a canard or a trailing-edge flap showed that, at approximately the design Mach number the cambered configuration developed a somewhat higher value than the trailing-edge flap configuration but a lower value than the canard configuration.

Adams, Gaynor J.↗

NASA NTRS1960

Performance, Stability, and Control Investigation at Mach Numbers from 0.60 to 1.05 of a Model of the "Swallow" with Outer Wing Panels Swept 75 degree with and without Power Simulations

An investigation of the performance, stability, and control characteristics of a variable-sweep arrow-wing model with the outer wing panels swept 75 deg. has been conducted in the Langley 16-foot transonic tunnel. Four outboard engines located above and below the wing provided propulsive thrust, and, by deflecting in the pitch direction and rotating in the lateral plane, also produced control forces. The engine nacelles incorporated swept lateral and vertical fins for aerodynamic stability and control. Jet-off data were obtained with flow-through nacelles, simulating inlet flow; jet thrust and hot-jet interference effects were obtained with faired-nose nacelles housing hydrogen peroxide gas generators. Six-component force and moment data were obtained at Mach numbers from 0.60 to 1.05 through a range of angles of attack and angles of side-slip. Control characteristics were obtained by deflecting the nacelle-fin combinations as elevators, rudders, and ailerons at several fixed angles for each control. The results indicate that the basic wing-body configuration becomes neutrally stable or unstable at a lift coefficient of 0.15; addition of nacelles with fins delayed instability to a lift coefficient of 0.30. Addition of nacelles to the wing-body configuration increased minimum drag from 0.0058 to 0.0100 at a Mach number of 0.60 and from 0.0080 to 0.0190 at a Mach number of 1.05 with corresponding reductions in maximum lift-drag ratio of 12 percent and 33 percent, respectively. The nacelle-fin combinations were ineffective as longitudinal controls but were adequate as directional and lateral controls. The model with nacelles and fins was directionally and laterally stable; the stability generally increased with increasing lift. Jet interference effects on stability and control characteristics were small but the adverse effects on drag were greater than would be expected for isolated nacelles.

Schmeer, James W.↗

NASA NTRS1961

A Flight Examination of Operating Problems of V/STOL Aircraft in STOL-Type Landing and Approach

A flight investigation has been conducted using a large twin-engine cargo aircraft to isolate the problems associated with operating propeller-driven aircraft in the STOL speed range where appreciable engine power is used to augment aerodynamic lift. The problems considered would also be representative of those of a large overloaded VTOL aircraft operating in an STOL manner with comparable thrust-to-weight ratios. The study showed that operation at low approach speeds was compromised by the necessity of maintaining high thrust to generate high lift and yet achieving the low lift-drag ratios needed for steep descents. The useable range of airspeed and flight path angle was limited by the pilot's demand for a positive climb margin at the approach speed, a suitable stall margin, and a control and/or performance margin for one engine inoperative. The optimum approach angle over an obstacle was found to be a compromise between obtaining the shortest air distance and the lowest touchdown velocity. In order to realize the greatest low-speed potential from STOL designs, the stability and control characteristics must be satisfactory.

Innis, Robert C.↗

NASA NTRS1990

Advanced recovery systems wind tunnel test report

Pioneer Aerospace Corporation (PAC) conducted parafoil wind tunnel testing in the NASA-Ames 80 by 120 test sections of the National Full-Scale Aerodynamic Complex, Moffett Field, CA. The investigation was conducted to determine the aerodynamic characteristics of two scale ram air wings in support of air drop testing and full scale development of Advanced Recovery Systems for the Next Generation Space Transportation System. Two models were tested during this investigation. Both the primary test article, a 1/9 geometric scale model with wing area of 1200 square feet and secondary test article, a 1/36 geometric scale model with wing area of 300 square feet, had an aspect ratio of 3. The test results show that both models were statically stable about a model reference point at angles of attack from 2 to 10 degrees. The maximum lift-drag ratio varied between 2.9 and 2.4 for increasing wing loading.

Geiger, R. H.↗

NASA NTRS1979

Effect of nose bluntness and afterbody shape on aerodynamic characteristics of a monoplanar missile concept with bodies of circular and elliptical cross sections at a Mach number of 2.50

The tests were performed at a Mach number of 2.50 and at angles of attack from about -4 deg to 32 deg. The results indicate that increasing nose bluntness increases zero lift drag and decreases both the maximum lift-drag ratio and the level of directional stability. The center of pressure generally moves forward with increasing nose size; however, small nose radii on the modified elliptical configurations move the center of pressure rearward. The circular bodied configurations exhibit the greatest longitudinal stability and the least directional stability. Concepts with the variable geometry afterbody contour display the most directional stability and the greatest zero lift drag.

Graves, E. B.↗

NASA NTRS1980

Conceptual study of an advanced supersonic technology transport (AST-107) for transpacific range using low-bypass-ratio turbofan engines

An advanced supersonic technology configuration concept designated the AST-107, using a low bypass ratio turbofan engine, is described and analyzed. The aircraft had provisions for 273 passengers arranged five abreast. The cruise Mach number was 2.62. The mission range for the AST-107 was 8.48 Mm (4576 n.mi.) and an average lift drag ratio of 9.15 during cruise was achieved. The available lateral control was not sufficient for the required 15.4 m/s (30 kt) crosswind landing condition, and a crosswind landing gear or a significant reduction in dihedral effect would be necessary to meet this requirement. The lowest computed noise levels, including a mechanical suppressor noise reduction of 3 EPNdB at the flyover and sideline monitoring stations, were 110.3 EPNdB (sideline noise), 113.1 EPNdB (centerline noise) and 110.5 EPNdB (approach noise).

Morris, S. J., Jr.↗

NASA NTRS1939

Observations in Flight of the Region of Stalled Flow over the Blades of an Autogiro Rotor

The flow over the inner halves of the rotor blades on a Kellet YG-1B autogiro was investigated in flight by making camera records of the motion of silk streamers attached to the upper surfaces of the blades. These records were analyzed to determine the boundaries of the region within which the flow over the blade sections was stalled for various tip-speed ratios. For the sake of comparison, corresponding theoretical boundaries were obtained. Both the size of the stalled area and its rate of growth with increasing tip-speed ratio were found to be larger than the theory predicted, although experiment agreed with theory with regard to shape and general location of the stalled area. The stalled region may be an important factor in both the rotor lift-drag ratio and the blade flapping motion at the higher tip-speed ratios. The method of study used in this paper should be useful in further studies of the problem, including the reduction of the size of the region.

Bailey, F J , Jr↗

NASA NTRS1992

Supersonic HLFC: Potential benefits and technology development requirements

The viewgraphs and discussion of the potential benefits and technology development requirements of supersonic Hybrid Laminar Flow Control (HLFC) are reviewed. For the last three years, Boeing has performed studies on the application of laminar flow control to High Speed Civil Transport (HSCT) configurations. Large potential net benefits were identified for laminar flow control, even after accounting for the significant implementation penalties. However, the technical risks are high at this time, and an early, aggressive technology development program is required if laminar flow control (LFC) is to be incorporated in a year 2005 HSCT program. The benefits and required development effort are addressed. Of all the aerodynamic advances that are being considered for the HSCT, LFC has the largest potential for improving the supersonic lift drag ratio of a given configuration. The work accomplished to date, sponsored by NASA Langley, is summarized. This work includes studies on HLFC application to HSCT, cruise HLFC/low speed boundary layer control (BLC) compatibility, and impact of M 0.9 HLFC and high lift BLC. Requirements for production committments are listed. The need for a HSCT-HLFC Risk Reduction Program is addressed. Supersonic HLFC development planning and future applications are addressed. Laminar flow/high-lift integration is addressed.

Neumann, Frank↗

NASA NTRS1959

Static Stability Characteristics of a Series of Hypersonic Boost-Glide Configurations at Mach Numbers of 1.41 and 2.01

An investigation of the static stability characteristics of several hypersonic boost-glide configurations has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41 and 2.01 (with Reynolds numbers per foot of 2.90 x 10(exp 6) and 2.41 x 10(exp 6) respectively). This series of configurations consisted of a cone, with and without cruciform fins, a trihedron, two low-aspect-ratio delta wings that differed primarily in cross-sectional shape, and two wing-body configurations. All configurations indicated reasonably linear pitching-, yawing-, and rolling-moment characteristics for angles of attack to at least 12 deg. The maximum lift-drag ratio for the zero-thrust condition (base drag included) was about 3 for the delta-wing configurations and about 4 for the wing-body configurations.

Foster, Gerald V.↗

NASA NTRS1954

Effects of Sweep and Thickness on the Static Longitudinal Aerodynamic Characteristics of a Series of Thin, Low-aspect-ratio, Highly Tapered Wings at Transonic Speeds : Transonic-bump Method

An investigation by the transonic-bump technique of the static longitudinal aerodynamic characteristics of a series of thin, low-aspect-ratio, highly tapered wings has been made in the Langley high-speed 7- by 10-foot tunnel. The Mach number range extended from about 0.60 to 1.18, with corresponding Reynolds numbers ranging from about 0.75 x 10(6) to 0.95 x 10(6). The angle of attack range was from -10 degrees to approximately 32 degrees.The effects on drag and lift-drag ratio of a variation in sweep angle from -14.03 degrees to 45 degrees with respect to the quarter-chord line for wings of 3-percent-chord thickness was found to be small in comparison to the effects of a variation in thickness from 2 percent chord to 4.5 percent chord for wings with 14.03 degree sweepback. For the range of variables considered, variations in plan form were considerably more important with regard to longitudinal stability characteristics than the variations in thickness. For the series of basic wings having an aspect ratio of 4, the most hearly linear pitching-moment characteristics were obtained with 26.57 degree of sweepback of the quarter-chord line. However, for the modified series of wings (obtained by clipping the tips of the original wings parallel to the plane of symmetry to give an aspect ratio of 3 and a taper ratio of 0.143), the most nearly linear pitching-moment characteristics were obtained with 36.87 degrees of sweepback. By decreasing the thickness-to-chord ratios from 0.03 to 0.02, a large increase in lift-curve slope was obtained for both the basic and modified wings. All of the wings of both series had fairly large inward shifts of the lateral center-of-pressure location (indicative of tip stalling) with increasing lift coefficient, except those wings having minimum sweepback angles.

Fournier, Paul G↗

NASA NTRS1974

Comparisons of two-dimensional shock-expansion theory with experimental aerodynamic data for delta-planform wings at high supersonic speeds

An investigation has been conducted to explore the potential for optimizing airfoil shape at high supersonic speeds by utilizing the two-dimensional shock-expansion method. Theoretical and experimental force and moment coefficients are compared for four delta-planform semispan wings having a leading-edge sweep angle of 65 deg and incorporating modified diamond airfoils with a thickness-chord ratio of 0.06. The wings differ only in airfoil maximum-thickness position and camber. The experimental data are obtained at Mach numbers of 3.95 and 4.63 and at a Reynolds number of 9.84 million per meter. A relatively simple method is developed for predicting, in terms of lift-drag ratio, the optimum modified diamond airfoil at high supersonic and hypersonic speeds.

Jernell, L. S.↗

NASA NTRS1961

Dynamic Stability and Control Problems of Piloted Reentry from Lunar Missions

A fixed-base simulator investigation has been made of stability and control problems during piloted reentry from lunar missions. Reentries were made within constraints of acceleration and skipping, in which the pilot was given simulated navigation tasks of altitude and heading angle commands. Vehicles considered included a blunt-face, high-drag capsule, and a low-drag lifting cone, each of which had a trim lift-drag ratio of 0.5. With the provision of three-axis automatic damping, both vehicles were easily controlled through reentry after a brief pilot-training period. With all dampers out, safe reentries could be made and both vehicles were rated satisfactory for emergency operation. In damper-failure conditions resulting in inadequate Dutch roll damping, the lifting-cone vehicle exhibited control problems due to excessive dihedral effect and oscillatory acceleration effects.

Moul, Martin T.↗