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NASA NTRSDate not supplied

Flow Physics Study of Sweeping Jet Actuation on an NACA 0015 Swept Wing Configuration

Results from a recently completed active flow control (AFC) experiment on a low aspect ratio, AR = 4.35, 30 degree swept, semispan wing with a 30% chord trailing edge flap are presented. The constant chord NACA 0015 wing is configured with sweeping jet actuators at the flap shoulder to facilitate our study of three-dimensional, swept wing separated and AFC controlled flowfields on the flap. Two flap deflection angles (delta_f = 0 degrees and 20 degrees) are examined. Oil and tuft flow visualization data, steady and unsteady pressure data, and particle image velocimetry (PIV) complement the balance lift and drag data. These results are used to examine smooth body separation, fixed separation, and the AFC-controlled flap flowfield downstream of a sweeping jet actuator. A momentum coefficient of 1.1% reattaches the flow to the flap surface when flap delta_f = 20 degrees increasing lift by approximately 0.26. This fundamental study is a first step in a research effort focused on improving the efficiency of AFC systems for swept wing applications.

active flow control↗
NASA NTRS1977

Comparison of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an analysis code

Comparisons of experimental and theoretical drag characteristics for a 10-percent-thick supercritical airfoil using a new version of an advanced analysis code. Comparisons are made at near-design normal-force coefficients for Reynolds numbers from 2 to 11 million. Comments are made concerning various input parameters to the code.

Harris, C. D.↗
NASA NTRS1941

A simplified theoretical method of determining the characteristics of a lifting rotor in forward flight

Theoretical derived expressions for the flapping, the thrust, the torque, and the profile drag-lift ratio of nonfeathering rotor with hinged, rectangular, linearly twisted blades are given as simple functions of the inflow velocity and the blade pitch. Representative values of the coefficients of each of the terms in these expressions are tabulated for a series of specified values of the tip-speed ratio. Analysis indicates that the tabulated values can be used to calculate, with reasonable accuracy, the characteristics of any rotor of conventional design.

Bailey, F J , Jr↗
NASA NTRS2005

Aerocapture Performance Analysis of A Venus Exploration Mission

A performance analysis of a Discovery Class Venus Exploration Mission in which aerocapture is used to capture a spacecraft into a 300km polar orbit for a two year science mission has been conducted to quantify its performance. A preliminary performance assessment determined that a high heritage 70 sphere-cone rigid aeroshell with a 0.25 lift to drag ratio has adequate control authority to provide an entry flight path angle corridor large enough for the mission s aerocapture maneuver. A 114 kilograms per square meter ballistic coefficient reference vehicle was developed from the science requirements and the preliminary assessment s heating indicators and deceleration loads. Performance analyses were conducted for the reference vehicle and for sensitivity studies on vehicle ballistic coefficient and maximum bank rate. The performance analyses used a high fidelity flight simulation within a Monte Carlo executive to define the aerocapture heating environment and deceleration loads and to determine mission success statistics. The simulation utilized the Program to Optimize Simulated Trajectories (POST) that was modified to include Venus specific atmospheric and planet models, aerodynamic characteristics, and interplanetary trajectory models. In addition to Venus specific models, an autonomous guidance system, HYPAS, and a pseudo flight controller were incorporated in the simulation. The Monte Carlo analyses incorporated a reference set of approach trajectory delivery errors, aerodynamic uncertainties, and atmospheric density variations. The reference performance analysis determined the reference vehicle achieves 100% successful capture and has a 99.87% probability of attaining the science orbit with a 90 meters per second delta V budget for post aerocapture orbital adjustments. A ballistic coefficient trade study conducted with reference uncertainties determined that the 0.25 L/D vehicle can achieve 100% successful capture with a ballistic coefficient of 228 kilograms per square meter and that the increased ballistic coefficient increases post aerocapture V budget to 134 meters per second for a 99.87% probability of attaining the science orbit. A trade study on vehicle bank rate determined that the 0.25 L/D vehicle can achieve 100% successful capture when the maximum bank rate is decreased from 30 deg/s to 20 deg/s. The decreased bank rate increases post aerocapture delta V budget to 102 meters per second for a 99.87% probability of attaining the science orbit.

Starr, Brett R.↗
NASA NTRS2006

Asymptotic Parachute Performance Sensitivity

In 2010, the Mars Science Laboratory mission will pioneer the next generation of robotic Entry, Descent, and Landing systems by delivering the largest and most capable rover to date to the surface of Mars. In addition to landing more mass than any other mission to Mars, Mars Science Laboratory will also provide scientists with unprecedented access to regions of Mars that have been previously unreachable. By providing an Entry, Descent, and Landing system capable of landing at altitudes as high as 2 km above the reference gravitational equipotential surface, or areoid, as defined by the Mars Orbiting Laser Altimeter program, Mars Science Laboratory will demonstrate sufficient performance to land on 83% of the planet s surface. By contrast, the highest altitude landing to date on Mars has been the Mars Exploration Rover at 1.3 km below the areoid. The coupling of this improved altitude performance with latitude limits as large as 60 degrees off of the equator and a precise delivery to within 10 km of a surface target, will allow the science community to select the Mars Science Laboratory landing site from thousands of scientifically interesting possibilities. In meeting these requirements, Mars Science Laboratory is extending the limits of the Entry, Descent, and Landing technologies qualified by the Mars Viking, Mars Pathfinder, and Mars Exploration Rover missions. Specifically, the drag deceleration provided by a Viking-heritage 16.15 m supersonic Disk-Gap-Band parachute in the thin atmosphere of Mars is insufficient, at the altitudes and ballistic coefficients under consideration by the Mars Science Laboratory project, to maintain necessary altitude performance and timeline margin. This paper defines and discusses the asymptotic parachute performance observed in Monte Carlo simulation and performance analysis and its effect on the Mars Science Laboratory Entry, Descent, and Landing architecture.

Way, David W.↗
NASA NTRS1972

Atmospheric Flight Mechanics Conference, 2nd, Palo Alto and Moffett Field, Calif., September 11-13, 1972, Informal Papers.

Limitations in the acquisition of nonlinear aerodynamic coefficients from free-oscillation data by means of the Chapman-Kirk technique, SAM-D control test vehicle trajectory plannning and flight test analysis, and determination of aerodynamic drag from radar data are among the topics covered in papers concerned with atmospheric flight mechanics. Other areas covered include fixed and rotary-wing aircraft, ordnance and reentry vehicles, and analysis and measurement techniques. Individual items are announced in this issue.

Source record↗
NASA NTRS1975

Aerodynamic Characteristics of a 14-Percent-Thick NASA Supercritical Airfoil Designed for a Normal-Force Coefficient of 0.7

This report documents the experimental aerodynamic characteristics of a 14 percent thick supercritical airfoil based on an off design sonic pressure plateau criterion. The design normal force coefficient was 0.7. The results are compared with those of the family related 10 percent thick supercritical airfoil 33. Comparisons are also made between experimental and theoretical characteristics and composite drag rise characteristics derived for a full scale Reynolds number of 40 million.

Harris, C. D.↗
NASA NTRS2011

Advanced High-Temperature Flexible TPS for Inflatable Aerodynamic Decelerators

Typical entry vehicle aeroshells are limited in size by the launch vehicle shroud. Inflatable aerodynamic decelerators allow larger aeroshell diameters for entry vehicles because they are not constrained to the launch vehicle shroud diameter. During launch, the hypersonic inflatable aerodynamic decelerator (HIAD) is packed in a stowed configuration. Prior to atmospheric entry, the HIAD is deployed to produce a drag device many times larger than the launch shroud diameter. The large surface area of the inflatable aeroshell provides deceleration of high-mass entry vehicles at relatively low ballistic coefficients. Even for these low ballistic coefficients there is still appreciable heating, requiring the HIAD to employ a thermal protection system (TPS). This TPS must be capable of surviving the heat pulse, and the rigors of fabrication handling, high density packing, deployment, and aerodynamic loading. This paper provides a comprehensive overview of flexible TPS tests and results, conducted over the last three years. This paper also includes an overview of each test facility, the general approach for testing flexible TPS, the thermal analysis methodology and results, and a comparison with 8-foot High Temperature Tunnel, Laser-Hardened Materials Evaluation Laboratory, and Panel Test Facility test data. Results are presented for a baseline TPS layup that can withstand a 20 W/cm2 heat flux, silicon carbide (SiC) based TPS layup, and polyimide insulator TPS layup. Recent work has focused on developing material layups expected to survive heat flux loads up to 50 W/cm2 (which is adequate for many potential applications), future work will consider concepts capable of withstanding more than 100 W/cm2 incident radiant heat flux. This paper provides an overview of the experimental setup, material layup configurations, facility conditions, and planned future flexible TPS activities.

DelCorso, Joseph A.↗
NASA NTRS2015

Adaptable, Deployable Entry and Placement Technology (ADEPT) Overview of FY15 Accomplishments

ADEPT is an atmospheric entry architecture for missions to most planetary bodies with atmospheres: Current Technology development project funded under STMD Game Changing Development Program (FY12 start); stowed inside the launch vehicle shroud and deployed in space prior to entry; low ballistic coefficient (less than 50 kilograms per square meter) provides a benign deceleration and thermal environment to the payload; High-temperature ribs support three dimensional woven carbon fabric to generate drag and withstand high heating.

technology↗
NASA NTRS1989

Lifetime predictions for the Solar Maximum Mission (SMM) and San Marco spacecraft

Lifetime prediction techniques developed by the Goddard Space Flight Center (GSFC) Flight Dynamics Division (FDD) are described. These techniques were developed to predict the Solar Maximum Mission (SMM) spacecraft orbit, which is decaying due to atmospheric drag, with reentry predicted to occur before the end of 1989. Lifetime predictions were also performed for the Long Duration Exposure Facility (LDEF), which was deployed on the 1984 SMM repair mission and is scheduled for retrieval on another Space Transportation System (STS) mission later this year. Concepts used in the lifetime predictions were tested on the San Marco spacecraft, which reentered the Earth's atmosphere on December 6, 1988. Ephemerides predicting the orbit evolution of the San Marco spacecraft until reentry were generated over the final 90 days of the mission when the altitude was less than 380 kilometers. The errors in the predicted ephemerides are due to errors in the prediction of atmospheric density variations over the lifetime of the satellite. To model the time dependence of the atmospheric densities, predictions of the solar flux at the 10.7-centimeter wavelength were used in conjunction with Harris-Priester (HP) atmospheric density tables. Orbital state vectors, together with the spacecraft mass and area, are used as input to the Goddard Trajectory Determination System (GTDS). Propagations proceed in monthly segments, with the nominal atmospheric drag model scaled for each month according to the predicted monthly average value of F10.7. Calibration propagations are performed over a period of known orbital decay to obtain the effective ballistic coefficient. Progagations using plus or minus 2 sigma solar flux predictions are also generated to estimate the despersion in expected reentry dates. Definitive orbits are compared with these predictions as time expases. As updated vectors are received, these are also propagated to reentryto continually update the lifetime predictions.

Smith, E. A.↗
NASA NTRS1923

Tests on an airplane model, AEG D I of the Allgemeine Elektricitats Gesellschft, A-G, airplane construction section conducted at the Gottingen Model Testing Laboratory for Aerodynamics

Tests were carried out in the small wind tunnel of the Gottingen establishment on a complete model of the AEG D I airplane. The agreement between the model and the complete airplane applies particularly to the wings, which have ribs cut out of sheet metal and built up in exactly the same manner as in the actual airplane. Various series of tests were carried out with this model in which one or the other of the control surfaces were adjusted to various angles, while the others remained in their neutral positions. During the first three series of tests, the stabilizer was set at a positive angle of 3 degrees, 45' relative to the axis of the engine crankshaft, after which further tests at a 6 degree 30' we made. Finally, the model was tested with the tail group removed. With the elevators set in the prescribed positions, the lift, the drag, and the moments about an axis passing through the center of gravity and perpendicular to the plane of symmetry were measured. All three sets of readings are given as absolute coefficients. Where one of the other control surfaces was deflected from its normal position, the moment produced by that adjustment of the surface was also measured.

Munk, Max↗
NASA NTRS1994

Two-dimensional converging-diverging rippled nozzles at transonic speeds

An experimental investigation was performed in the Langley 16-Foot Transonic tunnel to determine the effects of external and internal flap rippling on the aerodynamics of a nonaxisymmetric nozzle. Data were obtained at several Mach numbers from static conditions to 1.2 over a range of nozzle pressure ratios. Nozzles with chordal boattail angles of 10, 20, and 30 degrees, with and without surface rippling, were tested. No effect on discharge coefficient due to surface rippling was observed. Internal thrust losses due to surface rippling were measured and attributed to a combination of additional internal skin friction and shock losses. External nozzle drag for the baseline configurations were generally less than that for the rippled configurations at all free-stream Mach numbers tested. The difference between the baseline and rippled nozzle drag levels generally increased with increasing boat tail angle. The thrust-minus-drag level for each rippled nozzle configuration was less than the equivalent baseline configuration for each Mach number at the design nozzle pressure ratio.

Carlson, John R.↗
NASA NTRSDate not supplied

Transonic Lift and Drag Predictions using Wall Modelled Large Eddy Simulations

Wall-modelled Large Eddy Simulations (WMLES) of the NASA Common Research Model (CRM) at transonic conditions and various angles of attacks leading up to and including shock-induced flow separation are performed using the LAVA computational framework. The simulations are shown to accurately predict the lift curve slope and the onset of separation characterized by the break in the pitching moment. Furthermore, careful assessment of skin-friction drag at cruise condition and its subsequent decrease with increasing angles of attack are shown to be in agreement with viscous sublayer resolving Reynolds Averaged Navier Stokes(RANS) simulations. The small differences between WMLES and RANS appear to be of the same order as differences seen between two RANS models at the cruise-point. Some sensitivity is observed to the coefficient used in the subgrid scale model, although this can be reconciled by noting the low chord- and shock-incidence Reynolds numbers (Rex≈106) seen in the outboard regions of the wing, along with the uncertainties associated with tripping and the numerical transition that occurs near the leading edge. Among the primary drawbacks seen in the WMLES predictions is the subdued increase in wave-drag with increasing angle-of-attack and the shock intensity when compared with experimental data. Preliminary assessment suggests that additional span- and stream-wise grid resolution is likely needed in mid- and out-board portions of the wing to better resolve the shock-induced separated flow and to further investigate prediction accuracy of unsteady temporal characteristics of the problem.

ARMD↗
NASA NTRS1977

Improvement of maneuver aerodynamics by spanwise blowing

Spanwise blowing was used to test a generalized wind-tunnel model to investigate component concepts in order to provide improved maneuver characteristics for advanced fighter aircraft. Primary emphasis was placed on performance, stability, and control at high angles of attack and subsonic speeds. Test data were obtained in the Langley high speed 7 by 10 foot tunnel at free stream Mach numbers up to 0.50 for a range of model angles of attack, jet momentum coefficients, and leading and trailing edge flap deflection angles. Spanwise blowing on a 44 deg swept trapezoidal wing resulted in leading edge vortex enhancement with subsequent large vortex induced lift increments and drag polar improvements at the higher angles of attack. Small deflections of a leading edge flap delayed these lift and drag benefits to higher angles of attack. In addition, blowing was more effective at higher Mach numbers. Spanwise blowing in conjunction with a deflected trailing edge flap resulted in lift and drag benefits that exceeded the summation of the effects of each high lift device acting alone. Asymmetric blowing was an effective lateral control device at the higher angles of attack.

Erickson, G. E.↗
NASA NTRS1931

Lift and drag characteristics of a cabin monoplane

The results of flight tests conducted by the NACA to determine the lift and drag characteristics of a full-scale airplane are given herein. A Fairchild FC-2W2 cabin monoplane having a Gottingen 387 wing section was used for the tests. The maximum lift coefficient for the airplane is compared with that obtained for the Gottingen 387 airfoil in recent tests in the Variable Density Tunnel. The maximum lift coefficient for the airplane was found to be 1.50 and that for the airfoil 1.56. Although the flight tests were confined chiefly to glides with the propeller locked horizontally, data obtained with the propeller operating at zero thrust for a few angles of attack are also included.

Thompson, F L↗
NASA NTRS1957

Determination of Longitudinal Stability and Control Characteristics from Free-Flight Model Tests with Results at Transonic Speeds for Three Airplane Configurations

A test technique and data analysis method has been developed for determining the longitudinal aerodynamic characteristics from free-flight tests of rocket-propelled models. The technique makes use of accelerometers and an angle-of-attack indicator to permit instantaneous measurements of lift, drag, and pitching moments. The data, obtained during transient oscillations resulting from control-surface disturbances, are analyzed by essentially nonlinear direct methods (such as cross plots of the variation of lift coefficient with angle of attack) and by linear indirect methods by using the equations of motion for a transient oscillation. The analysis procedure has been set forth in some detail and the feasibility of the method has been demonstrated by data measured through the transonic speed range on several airplane configurations. It was shown that the flight conditions and dynamic similitude factors for the tests described were reasonably close to typical full-scale airplane conditions.

Gillis, Clarence L↗