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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Methods for determining the internal thrust of scramjet engine modules from experimental data

Methods for calculating zero-fuel internal drag of scramjet engine modules from experimental measurements are presented. These methods include two control-volume approaches, and a pressure and skin-friction integration. The three calculation techniques are applied to experimental data taken during tests of a version of the NASA parametric scramjet. The methods agree to within seven percent of the mean value of zero-fuel internal drag even though several simplifying assumptions are made in the analysis. The mean zero-fuel internal drag coefficient for this particular engine is calculated to be 0.150. The zero-fuel internal drag coefficient when combined with the change in engine axial force with and without fuel defines the internal thrust of an engine.

Voland, Randall T.↗
NASA NTRS1993

Generic hypersonic vehicle performance model

An integrated computational model of a generic hypersonic vehicle was developed for the purpose of determining the vehicle's performance characteristics, which include the lift, drag, thrust, and moment acting on the vehicle at specified altitude, flight condition, and vehicular configuration. The lift, drag, thrust, and moment are developed for the body fixed coordinate system. These forces and moments arise from both aerodynamic and propulsive sources. SCRAMjet engine performance characteristics, such as fuel flow rate, can also be determined. The vehicle is assumed to be a lifting body with a single aerodynamic control surface. The body shape and control surface location are arbitrary and must be defined. The aerodynamics are calculated using either 2-dimensional Newtonian or modified Newtonian theory and approximate high-Mach-number Prandtl-Meyer expansion theory. Skin-friction drag was also accounted for. The skin-friction drag coefficient is a function of the freestream Mach number. The data for the skin-friction drag coefficient values were taken from NASA Technical Memorandum 102610. The modeling of the vehicle's SCRAMjet engine is based on quasi 1-dimensional gas dynamics for the engine diffuser, nozzle, and the combustor with heat addition. The engine has three variable inputs for control: the engine inlet diffuser area ratio, the total temperature rise through the combustor due to combustion of the fuel, and the engine internal expansion nozzle area ratio. The pressure distribution over the vehicle's lower aft body surface, which acts as an external nozzle, is calculated using a combination of quasi 1-dimensional gas dynamic theory and Newtonian or modified Newtonian theory. The exhaust plume shape is determined by matching the pressure inside the plume, calculated from the gas dynamic equations, with the freestream pressure, calculated from Newtonian or Modified Newtonian theory. In this manner, the pressure distribution along the vehicle after body expansion surface is then determined. The aerodynamic modeling, the engine modeling, and the exhaust plume analysis are described in more detail. A description of the computer code used to perform the above calculations is given and an input/output example is then given. The computer code is available on a Macintosh floppy disk.

Chavez, Frank R.↗
NASA NTRS1940

Wind-tunnel Investigation of Two Airfoils with 25-percent-chord Gwinn and Plain Flaps

Aerodynamic force tests of an NACA 23018 airfoil with a Gwinn flap having a chord 25 percent of the overall chord and of an NACA 23015 airfoil with a plain flap having a 25-percent chord were conducted to determine the relative merits of the Gwinn and the plain flaps. The tests indicated that, based on speed-range ratios, the plain flap was more effective than the Gwinn flap. At small flap deflections, the plain flap had lower drag coefficients at lift-coefficient values less than 0.70. For lift coefficients greater than 0.70, however, the Gwinn flap at all downward flap deflections had the lower drag coefficients.

Ames, Milton B , Jr↗
DOE OSTI2021

On the temperature and density dependence of dislocation drag from phonon wind

At extreme strain rates, where fast moving dislocations govern plastic deformation, anharmonic phonon scattering imparts a drag force on the dislocations. In this paper, we present calculations of the dislocation drag coefficients of aluminum and copper as functions of temperature and density. We discuss the sensitivity of the drag coefficients to changes in the third-order elastic constants with temperature and density.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗
NASA NTRS1950

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

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.

Gillespie, Warren, Jr.↗
NASA NTRS1955

Aerodynamic Characteristics of a 0.04956-Scale Model of the Convair F-102A Airplane at Transonic Speeds

Tests have been conducted in the Langley 8-foot transonic tunnel on a 0.04956-scale model of the Convair F-102A airplane which employed an indented and extended fuselage, cambered wing leading edges, and deflected wing tips. Force and moment characteristics were obtained for Mach numbers from 0.60 to 1.135 at angles of attack up to 20 . In addition, tests were made over a limited angle-of-attack range to determine the effects of the cambered leading edges, deflected tips, and a nose section with a smooth area distribution. Fuselage modifications employed on the F-102A were responsible for a 25.percent reduction in the minimum drag-coefficient rise between the Mach numbers of 0.85 and 1.075 when compared with that for the earlier versions of the F-102. Although the wing modifications increased the F-102A subsonic minimum drag-coefficient level approximately 0.0020, they produced large decreases in drag at lifting conditions over that for the original (plane-wing) F-102. The F-102A had 15 to 25 percent higher maximum lift-drag ratios than did the original F-102. The F-102A had about 15 percent lower maximum lift-drag ratios at Mach numbers below 0.95 and slightly higher maximum lift-drag ratios at supersonic speeds when compared with those ratios for sn earlier modified-wing version of the F-102. Chordwise wing fences which provided suitable longitudinal stability for the original F-102 were not adequate for the cambered-wing F-102A The pitching-moment curves indicated a region of near neutral stability with possible pitch-up tendencies for the F-102A at high subsonic Mach numbers for lift coefficients between about 0.4 and 0.5.

Tempelmeyer, Kenneth E.↗
NASA NTRS1996

Rarefaction effects on Galileo probe aerodynamics

Solutions of aerodynamic characteristics are presented for the Galileo Probe entering Jupiter's hydrogen-helium atmosphere at a nominal relative velocity of 47.4 km/s. Focus is on predicting the aerodynamic drag coefficient during the transitional flow regime using the direct simulation Monte Carlo (DSMC) method. Accuracy of the probe's drag coefficient directly impacts the inferred atmospheric properties that are being extracted from the deceleration measurements made by onboard accelerometers as part of the Atmospheric Structure Experiment. The range of rarefaction considered in the present study extends from the free molecular limit to continuum conditions. Comparisons made with previous calculations and experimental measurements show the present results for drag to merge well with Navier-Stokes and experimental results for the least rarefied conditions considered.

Moss, James N.↗
NASA NTRS2018

Multidisciplinary Design Optimization and Cruise Mach Number Study of Truss-Braced Wing Aircraft

The Subsonic Ultra Green Aircraft Research (SUGAR) Phase III was led by Dr. Rakesh K. Kapania and Dr. Joseph A. Schetz at the Multidisciplinary Analysis and Design Center for Advanced Vehicles, Department of Aerospace and Ocean Engineering, Virginia Tech, Blacksburg VA. The research was performed from December 2014 to December 2015. Three major areas were investigated: Multidisciplinary Design Optimization (MDO) studies of truss braced wing (TBW) and strut braced wing (SBW) vehicles at cruise Mach numbers of 0.7 and 0.8 for a flight mission similar to current market single aisle configurations. The performance and the characteristics of the optimized vehicles were compared to the SUGAR Phase II TBW vehicle. These results were obtained without applying any of the extended transonic aerodynamic and aeroelastic tools that will be discussed later. It was found that the cruise Mach number has a large effect on the best “truss” configuration. At Mach 0.7, an SBW has a better fuel consumption and better take-off gross weight (TOGW). However, at Mach 0.8, the TBW is superior because the jury strut aids in satisfying the flutter constraint; Two-dimensional, steady, transonic aerodynamic analysis of the Boeing Airfoil J (BACJ) airfoil was performed for a range of thickness ratios, Mach numbers and lift coefficients. Reynolds-averaged Navier-Stokes (RANS) equations were solved to obtain the lift-curve slope, wave drag coefficient, the location of the center of pressure and to predict the separation at the trailing edge, which may lead to buffeting. One of the goals was to develop a database of lift-curve slope and the location of center of pressure, which could be used in a transonic aeroelastic analysis. Another goal was to compare the wave drag coefficients to those predicted by Lock’s fourth-power law and also to compare the transonic effects obtained from RANS simulations to those predicted by the Korn equations. A third goal was to develop a buffet boundary, which can be integrated into the MDO framework to prevent the optimized designs from probable buffeting; A state-space transonic aeroelastic analysis tool was developed, which can incorporate the nonlinear transonic effects in the unsteady aerodynamics but is yet computationally cheap when used within the VT MDO framework. The aeroelastic analysis uses Leishman- Beddoes (LB) indicial functions, which generated a state-space representation of the aeroelastic system. The indicial functions allow the incorporation of data for steady lift-curve slope and location of the center of pressure. Thus, the steady transonic effects are included, and the unsteady aerodynamic responses are a linearization about the steady results. The aeroelastic approach discretizes the wing into numerous strips, which results in a large eigenvalue problem as each strip has eight augmented aerodynamic states as per the LB theory. Thus, to reduce the computation expense, a reduced order model (ROM) was developed. The approach was validated using a few examples.

Kapania, Rakesh K.↗
NASA NTRS1974

Transonic off-design drag and performance of an axisymmetric inlet with 40 percent internal contraction on design

An experimental investigation determined the drag and pressure performance of an axisymmetric supersonic inlet when operated in the transonic speed range. The inlet configuration was derived from a Mach 2.5 mixed compression inlet design with assumed variable geometry. At typical engine airflows the drag coefficient varied from 0.057 to 0.192 when the Mach number changed from 0.80 to 1.27. The presence of a wing simulator resulted in a sizable increase in total drag at Mach 1.2. This interference drag, which is roughly a 0.1 increase in drag coefficient, originates equally from an increase in both additive and cowl pressure drag.

Woollett, R. R.↗
NASA NTRS1956

The Aerodynamic Characteristics in Pitch of a 1/15-Scale Model of the Grumman F11F-1 Airplane at Mach Numbers of 1.41, 1.61, and 2.01, TED No. NACA DE 390

Tests have been made in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41, 1.61, and 2.01 to determine the static longitudinal stability and control characteristics of various arrangements of the Grumman F11F-1 airplane. Tests were made of the complete model and various combinations of its component parts and, in addition, the effects of various body modifications, a revised vertical tail, and wing fences on the longitudinal characteristics were determined. The results indicate that for a horizontal-tail incidence of -10 deg the trim lift coefficient varied from 0.29 at a Mach number of 1.61 to 0.23 at a Mach number of 2.01 with a corresponding decrease in lift-drag trim from 3.72 to 3.15. Stick-position instability was indicated in the low-supersonic-speed range. A photographic-type nose modification resulted in slightly higher values of minimum drag coefficient but did not significantly affect the static stability or lift-curve slope. The minimum drag coefficient for the complete model with the production nose remained essentially constant at 0.047 throughout the Mach number range investigated.

Driver, Cornelius↗
NASA NTRS1980

Wind-tunnel/flight correlation study of aerodynamic characteristics of a large flexible supersonic cruise airplane (XB-70-1). 3: A comparison between characteristics predicted from wind-tunnel measurements and those measured in flight

A program was undertaken by NASA to evaluate the accuracy of a method for predicting the aerodynamic characteristics of large supersonic cruise airplanes. This program compared predicted and flight-measured lift, drag, angle of attack, and control surface deflection for the XB-70-1 airplane for 14 flight conditions with a Mach number range from 0.76 to 2.56. The predictions were derived from the wind-tunnel test data of a 0.03-scale model of the XB-70-1 airplane fabricated to represent the aeroelastically deformed shape at a 2.5 Mach number cruise condition. Corrections for shape variations at the other Mach numbers were included in the prediction. For most cases, differences between predicted and measured values were within the accuracy of the comparison. However, there were significant differences at transonic Mach numbers. At a Mach number of 1.06 differences were as large as 27 percent in the drag coefficients and 20 deg in the elevator deflections. A brief analysis indicated that a significant part of the difference between drag coefficients was due to the incorrect prediction of the control surface deflection required to trim the airplane.

Arnaiz, H. H.↗
NASA NTRS1984

Experimental investigation of turbulent flow in smooth and longitudinal grooved tubes

Turbulent flow in tubes with and without longitudinal grooves is examined. The discovery of fine grooves forming a sort of streamline pattern on the body of sharks led to the expectation that the grooves on a surface reduce the momentum change, and thus the drag. To test this thesis, drag law, velocity profile and the profile of the velocity fluctuation were determined. Results show that for moderate Reynolds numbers the drag coefficient for grooved tubes is about 3 percent smaller than that of the smooth tubes. At higher Reynolds numbers, however, the drag coefficient for grooved tubes becomes larger than that for smooth tubes. No significant differences in the velocity profiles between grooved tubes and smooth tubes are found.

Nitschke, P.↗
NASA NTRS2005

Should We Believe Atmospheric Temperatures Measured by Entry Accelerometers Traveling at "Slow" Near-Sonic Speeds?

Mars Pathfinder's Accelerometer instrument measured an unexpected and large temperature inversion between 10 and 20 kilometer altitude. Other instruments have failed to detect similar temperature inversions. I test whether this inversion is real or not by examining what changes have to be made to the assumptions in the accelerometer data processing to obtain a more "expected" temperature profile. Changes in derived temperature of up to 30K, or 15%, are necessary, which correspond to changes in derived density of up to 25% and changes in derived pressure of up to 10%. If the drag coefficient is changed to satisfy this, then instead of decreasing from 1.6 to 1.4 from 20 kilometers to 10 kilometers, the drag coefficient must increase from 1.6 to 1.8 instead. If winds are invoked, then speeds of 60 meters per second are necessary, four times greater than those predicted. Refinements to the equation of hydrostatic equilibrium modify the temperature profile by an order of magnitude less than the desired amount. Unrealistically large instrument drifts of 0.5-1.0 meters per square second are needed to adjust the temperature profile as desired. However, rotational contributions to the accelerations may have the necessary magnitude and direction to make this correction. Determining whether this hypothesis is true will require further study of the rigid body equations of motion, with detailed knowledge of the positions of all six accelerometers. The paradox concerning this inversion is not yet resolved. It is important to resolve it because the paradox has some startling implications. At one extreme, are temperature profiles derived from accelerometers inherently inaccurate by 20K or more? At the other extreme, are RS temperature profiles inaccurate by this same amount?

Withers, Paul↗
NASA NTRS1959

Effects of Outboard Thickened and Blunted Leading Edges on the Wave Drag of a 45 Degree Swept-Wing and Body Combination

An investigation to evaluate the effects of thickened and blunted leading-edge modifications on the wave drag of a swept wing has been made at Mach numbers from 0.65 to 2.20 and at a Reynolds number of 2,580,000 based on the mean aerodynamic chord of the basic wing. Two leading-edge designs were investigated and they are referred to as the thickened and the blunted modifications although both sections had equally large leading-edge radii. The thickened leading edge was formed by increasing the thickness over the forward 40 percent of the basic wing section. The blunted modification was formed by reducing the wing chords about 1 percent and by increasing the section thickness slightly over the forward 6 percent of the basic section in a manner to keep the wing sweep and volume essentially equal to the respective values for the basic wing. The basic wing had an aspect ratio of 3, a leading-edge sweep of 45 deg., a taper ratio of 0.4, and NACA 64AO06 sections perpendicular to a line swept back 39.45 deg., the quarter-chord line of these sections. Test results indicated that the thickened modification resulted in an increase in zero-lift drag coefficient of from 0.0040 to 0.0060 over values for the basic model at Mach numbers at which the wing leading edge was sonic or supersonic. Although drag coefficients of both the basic and thickened models were reduced at all test Mach numbers by body indentations designed for the range of Mach numbers from 1.00 to 2.00, the greater drag of the thickened model relative to that of the basic model was not reduced. The blunted model, however, had less than one quarter of the drag penalty of the thickened model relative to the basic model at supersonic leading-edge conditions (M greater or equal to root-2).

Holdaway, George H.↗
NASA NTRS1986

Advanced natural laminar flow airfoil with high lift to drag ratio

An experimental verification of a high performance natural laminar flow (NLF) airfoil for low speed and high Reynolds number applications was completed in the Langley Low Turbulence Pressure Tunnel (LTPT). Theoretical development allowed for the achievement of 0.70 chord laminar flow on both surfaces by the use of accelerated flow as long as tunnel turbulence did not cause upstream movement of transition with increasing chord Reynolds number. With such a rearward pressure recovery, a concave type deceleration was implemented. Two-dimensional theoretical analysis indicated that a minimum profile drag coefficient of 0.0026 was possible with the desired laminar flow at the design condition. With the three-foot chord two-dimensional model constructed for the LTPT experiment, a minimum profile drag coefficient of 0.0027 was measured at c sub l = 0.41 and Re sub c = 10 x 10 to the 6th power. The low drag bucket was shifted over a considerably large c sub l range by the use of the 12.5 percent chord trailing edge flap. A two-dimensional lift to drag ratio (L/D) was 245. Surprisingly high c sub l max values were obtained for an airfoil of this type. A 0.20 chort split flap with 60 deg deflection was also implemented to verify the airfoil's lift capabilities. A maximum lift coefficient of 2.70 was attained at Reynolds numbers of 3 and 6 million.

Viken, Jeffrey K.↗
NASA NTRS1977

Pressure and heating-rate distributions on a corrugated surface in a supersonic turbulent boundary layer

Drag and heating rates on wavy surfaces typical of current corrugated plate designs for thermal protection systems were determined experimentally. Pressure-distribution, heating-rate, and oil-flow tests were conducted in the Langley Unitary Plan wind tunnel at Mach numbers of 2.4 and 4.5 with the corrugated surface exposed to both thick and thin turbulent boundary layers. Tests were conducted with the corrugations at cross-flow angles from 0 deg to 90 deg to the flow. Results show that for cross-flow angles of 30 deg or less, the pressure drag coefficients are less than the local flat-plate skin-friction coefficients and are not significantly affected by Mach number, Reynolds number, or boundary-layer thickness over the ranges investigated. For cross-flow angles greater than 30 deg, the drag coefficients increase significantly with cross-flow angle and moderately with Reynolds number. Increasing the Mach number causes a significant reduction in the pressure drag. The average and peak heating penalties due to the corrugated surface are small for cross-flow angles of 10 deg or less but are significantly higher for the larger cross-flow angles.

Sawyer, J. W.↗
NASA NTRS2006

A Summary of the Experimental Results for a Generic Tractor-Trailer in the Ames Research Center 7- by 10-Foot and 12-Foot Wind Tunnels

Experimental measurements of a generic tractor-trailer were obtained in two wind tunnels at Ames Research Center. After a preliminary study at atmospheric conditions in the 7- by 10-Foot Wind Tunnel, additional testing was conducted at Reynolds numbers corresponding to full-scale highway speeds in the 12-Foot Pressure Wind Tunnel. To facilitate computational modeling, the 1:8-scale geometry, designated the Generic Conventional Model, included a simplified underbody and omitted many small-scale details. The measurements included overall and component forces and moments, static and dynamic surface pressures, and three-component particle image velocimetry. This summary report highlights the effects of numerous drag reduction concepts and provides details of the model installation in both wind tunnels. To provide a basis for comparison, the wind-averaged drag coefficient was tabulated for all configurations tested. Relative to the baseline configuration representative of a modern class-8 tractor-trailer, the most effective concepts were the trailer base flaps and trailer belly box providing a drag-coefficient reduction of 0.0855 and 0.0494, respectively. Trailer side skirts were less effective yielding a drag reduction of 0.0260. The database of this experimental effort is publicly available for further analysis.

Storms, Bruce L.↗
NASA NTRS1982

A model for unsteady transonic indicial responses

A method of characterizing the indicial curve for the response frequencies of flutter in the transonic regime by means of a time scale parameter is illustrated. Time linearization of the unsteady potential equation is assumed adequate and the perturbed potential about a steady flow is defined with appropriate boundary conditions. An indicial motion can then be obtained for a given mode, with the lift and drag coefficients which result from a step change in incidence or control surface deflection available to describe any kind of motion. The lift and drag coefficients are shown to result analytically by addition of the multiplicand involving the time scale parameter. If the coefficients are maximized, then at reduced frequencies only two parameters are necessary to form the indicial response curve.

Fung, K.-Y.↗