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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS2001

Solar Wind Influence on the Oxygen Content of Ion Outflow in the High-Altitude Polar Cap During Solar Minimum Conditions

We correlate solar wind and interplanetary magnetic field (IMF) properties with the properties of O(+) and H(+) during early 1996 (solar minimum) at altitudes between 5.5 and 8.9 R(sub E) geocentric using the Thermal Ion Dynamics Experiment (TIDE) on the Polar satellite. Throughout the high-altitude polar cap we observe H(+) to be more abundant than O(+). O(+) is found to be more abundant at lower latitudes when the solar wind speed is low (and Kp is low), while at higher solar wind speeds (and high Kp), O(+) is observed across most of the polar cap. The O(+) density and parallel flux are well organized by solar wind dynamic pressure, both increasing with solar wind dynamic pressure. Both the O(+) density and parallel flux have positive correlations with both V(sub SW)B(sub IMF) and E(sub SW). No correlation is found between O(+) density and IMF Bz, although a nonlinear relationship with IMF By is observed, possibly due to a strong linear correlation with the dynamic pressure. H(+) is not as highly correlated with solar wind and IMF parameters, although H(+) density and parallel flux are negatively correlated with IMF By and positively correlated with both V(sub SW)B(sub IMF) and E(sub SW). In this solar minimum data set, H(+) is dominant, so that contributions of this plasma to the plasma sheet would have very low O(+) to H(+) ratios.

Elliott, H. A.↗
NASA NTRS1989

Performance and test section flow characteristics of the National Full-Scale Aerodynamics Complex 40- by 80-foot wind tunnel

Results from the performance and test section flow calibration of the 40- by 80-Foot Wind Tunnel are presented. A flow calibration test was conducted in May and June 1987. The goal of the flow calibration test was to determine detailed spatial variations in the 40- by 80-ft test section flow quality throughout the tunnel operational envelope. Data were collected for test section speeds up to 300 knots and for air exchange rates of 0, 5, and 10 percent. The tunnel performance was also calibrated during the detailed mapping of the test section flow field. Experimental results presented indicate that the flow quality in the test section, with the exception of temperature, is relatively insensitive to the level of dynamic pressure and the air exchange rate. The dynamic pressure variation in the test section is within + or - 0.5 deg at all test section velocities. Cross-stream temperature gradients in the test section caused by the air exchange system were documented, and a correction method was established. Streamwise static pressure variation on the centerline is about 1 percent of test section dynamic pressure over 30 ft of the test section length.

Zell, Peter T.↗
NASA NTRS1991

Solar wind control of the magnetopause shape, location, and motion

A data set of 1821 magnetopause crossings was assembled. Separate fits to subsets of this data set determine the magnetopause location as a function of solar wind dynamic pressure and interplanetary magnetic field orientation. Solar wind dynamic pressure variations produce self-similar magnetopause motion on time scales of one hour or longer. In this paper, the pressure balance relationship between the solar wind dynamic pressure and the location of the subsolar magnetopause are verified. The relationship between the IMF Bz, region 1 Birkeland current strength, the position of the subsolar magnetopause, and the shape of the dayside magnetosphere is quantified. Cross sections of the dayside magnetopause in planes perpendicular to the earth-sun line are oblate.

Sibeck, D. G.↗
NASA NTRS2016

Experimental Verification of Buffet Calculation Procedure Using Unsteady PSP

Typically a limited number of dynamic pressure sensors are employed to determine the unsteady aerodynamic forces on large, slender aerospace structures. The estimated forces are known to be very sensitive to the number of the dynamic pressure sensors and the details of the integration scheme. This report describes a robust calculation procedure, based on frequency-specific correlation lengths, that is found to produce good estimation of fluctuating forces from a few dynamic pressure sensors. The validation test was conducted on a flat panel, placed on the floor of a wind tunnel, and was subjected to vortex shedding from a rectangular bluff-body. The panel was coated with fast response Pressure Sensitive Paint (PSP), which allowed time-resolved measurements of unsteady pressure fluctuations on a dense grid of spatial points. The first part of the report describes the detail procedure used to analyze the high-speed, PSP camera images. The procedure includes steps to reduce contamination by electronic shot noise, correction for spatial non-uniformities, and lamp brightness variation, and finally conversion of fluctuating light intensity to fluctuating pressure. The latter involved applying calibration constants from a few dynamic pressure sensors placed at selective points on the plate. Excellent comparison in the spectra, coherence and phase, calculated via PSP and dynamic pressure sensors validated the PSP processing steps. The second part of the report describes the buffet validation process, for which the first step was to use pressure histories from all PSP points to determine the "true" force fluctuations. In the next step only a selected number of pixels were chosen as "virtual sensors" and a correlation-length based buffet calculation procedure was applied to determine "modeled" force fluctuations. By progressively decreasing the number of virtual sensors it was observed that the present calculation procedure was able to make a close estimate of the "true" unsteady forces only from four sensors. It is believed that the present work provides the first validation of the buffet calculation procedure which has been used for the development of many space vehicles.

Buffet↗
NASA NTRS2015

In-Flight Aeroelastic Stability of the Thermal Protection System on the NASA HIAD, Part II: Nonlinear Theory and Extended Aerodynamics

Conical shell theory and a supersonic potential flow aerodynamic theory are used to study the nonlinear pressure buckling and aeroelastic limit cycle behavior of the thermal protection system for NASA's Hypersonic Inflatable Aerodynamic Decelerator. The structural model of the thermal protection system consists of an orthotropic conical shell of the Donnell type, resting on several circumferential elastic supports. Classical Piston Theory is used initially for the aerodynamic pressure, but was found to be insufficient at low supersonic Mach numbers. Transform methods are applied to the convected wave equation for potential flow, and a time-dependent aerodynamic pressure correction factor is obtained. The Lagrangian of the shell system is formulated in terms of the generalized coordinates for all displacements and the Rayleigh-Ritz method is used to derive the governing differential-algebraic equations of motion. Aeroelastic limit cycle oscillations and buckling deformations are calculated in the time domain using a Runge-Kutta method in MATLAB. Three conical shell geometries were considered in the present analysis: a 3-meter diameter 70 deg. cone, a 3.7-meter 70 deg. cone, and a 6-meter diameter 70 deg. cone. The 6-meter configuration was loaded statically and the results were compared with an experimental load test of a 6-meter HIAD. Though agreement between theoretical and experimental strains was poor, the circumferential wrinkling phenomena observed during the experiments was captured by the theory and axial deformations were qualitatively similar in shape. With Piston Theory aerodynamics, the nonlinear flutter dynamic pressures of the 3-meter configuration were in agreement with the values calculated using linear theory, and the limit cycle amplitudes were generally on the order of the shell thickness. The effect of axial tension was studied for this configuration, and increasing tension was found to decrease the limit cycle amplitudes when the circumferential elastic supports were neglected, but resulted in more complex behavior when the supports were included. The nominal flutter dynamic pressure of the 3.7-meter configuration was significantly lower than that of the 3-meter, and it was found that two sets of natural modes coalesce to flutter modes near the same dynamic pressure. This resulted in a significant drop in the limit cycle frequencies at higher dynamic pressures, where the flutter mode with the lower frequency becomes more critical. Pre-buckling pressure loads and the aerodynamic pressure correction factor were studied for all geometries, and these effects resulted in significantly lower flutter boundaries compared with Piston Theory alone. The maximum dynamic pressure predicted by aerodynamic simulations of a proposed 3.7-meter HIAD vehicle was still lower than any of the calculated flutter dynamic pressures, suggesting that aeroelastic effects for this vehicle are of little concern.

Goldman, Benjamin D.↗
NASA NTRS1990

The effectiveness of vane-aileron excitation in the experimental determination of flutter speed by parameter identification

The effectiveness of aerodynamic excitation is evaluated analytically in conjunction with the experimental determination of flutter dynamic pressure by parameter identification. Existing control surfaces were used, with an additional vane located at the wingtip. The equations leading to the identification of the equations of motion were reformulated to accommodate excitation forces of aerodynamic origin. The aerodynamic coefficients of the excitation forces do not need to be known since they are determined by the identification procedure. The 12 degree-of-freedom numerical example treated in this work revealed the best wingtip vane locations, and demonstrated the effectiveness of the aileron-vane excitation system. Results from simulated data gathered at much lower dynamic pressures (approximately half the value of flutter dynamic pressure) predicted flutter dynamic pressures with 2-percent errors.

Nissim, Eli↗
NASA NTRS1990

Effects of spoiler surfaces on the aeroelastic behavior of a low-aspect-ratio rectangular wing

An experimental research study to determine the effectiveness of spoiler surfaces in suppressing flutter onset for a low-aspect-ratio, rectangular wing was conducted in the Langley Transonic Dynamics Tunnel (TDT). The wing model used in this flutter test consisted of a rigid wing mounted to the wind-tunnel wall by a flexible, rectangular beam. The flexible beam was connected to the wing root and cantilever mounted to the wind-tunnel wall. The wing had a 1.5 aspect ratio based on wing semispan and a NACA 64A010 airfoil shape. The spoiler surfaces consisted of thin, rectangular aluminum plates that were vertically mounted to the wing surface. The spoiler surface geometry and location on the wing surface were varied to determine the effects of these parameters on the classical flutter of the wing model. Subsonically, the experiment showed that spoiler surfaces increased the flutter dynamic pressure with each successive increase in spoiler height or width. This subsonic increase in flutter dynamic pressure was approximately 15 percent for the maximum height spoiler configuration and for the maximum width spoiler configuration. At transonic Mach numbers, the flutter dynamic pressure conditions were increased even more substantially than at subsonic Mach numbers for some of the smaller spoiler surfaces. But greater than a certain spoiler size (in terms of either height or width) the spoilers forced a torsional instability in the transonic regime that was highly Mach number dependent. This detrimental torsional instability was found at dynamic pressures well below the expected flutter conditions. Variations in the spanwise location of the spoiler surfaces on the wing showed little effect on flutter. Flutter analysis was conducted for the basic configuration (clean wing with all spoiler surface mass properties included). The analysis correlated well with the clean wing experimental flutter results.

Cole, Stanley R.↗
NASA NTRS1979

Experimental study of acoustic loads on an upper-surface-blown STOL airplane configuration

Fluctuating pressure levels were measured on the flap and fuselage of an upper-surface-blown jet-flap airplane configuration in a wind tunnel. The model tested had turbofan engines with a bypass ratio of 3 and a thrust rating of 10 kN. Rectangular nozzles were mounted flush with the upper surface at 35 percent of the wing chord. Test parameters were flap deflection angle, jet impingement angle, angle of attack, free-stream velocity, spanwise location of the engine, and jet dynamic pressure. Load levels were high throughout the jet impingement region, with the highest levels (about 159 dB) occurring on the fuselage and near the knee of the flap. The magnitude of the forward-velocity effect appeared to depend upon the ratio of free-stream and jet velocities. Good agreement was obtained between fluctuating pressure spectra measured at jet dynamic pressures of 7 and 22 kPa when the spectra were scaled by nondimensional functions of dynamic pressure, velocity, and the empirical relationship between dynamic pressure and overall fluctuating pressure level.

Willis, C. M.↗
NASA NTRS1997

Experimental Investigation of Unsteady Shock Wave Turbulent Boundary Layer Interactions About a Blunt Fin

A series of experiments were performed to investigate the effects of Mach number variation on the characteristics of the unsteady shock wave/turbulent boundary layer interaction generated by a blunt fin. A single blunt fin hemicylindrical leading edge diameter size was used in all of the experiments which covered the Mach number range from 2.0 to 5.0. The measurements in this investigation included surface flow visualization, static and dynamic pressure measurements, both on centerline and off-centerline of the blunt fin axis. Surface flow visualization and static pressure measurements showed that the spatial extent of the shock wave/turbulent boundary layer interaction increased with increasing Mach number. The maximum static pressure, normalized by the incoming static pressure, measured at the peak location in the separated flow region ahead of the blunt fin was found to increase with increasing Mach number. The mean and standard deviations of the fluctuating pressure signals from the dynamic pressure transducers were found to collapse to self-similar distributions as a function of the distance perpendicular to the separation line. The standard deviation of the pressure signals showed initial peaked distribution, with the maximum standard deviation point corresponding to the location of the separation line at Mach number 3.0 to 5.0. At Mach 2.0 the maximum standard deviation point was found to occur significantly upstream of the separation line. The intermittency distributions of the separation shock wave motion were found to be self-similar profiles for all Mach numbers. The intermittent region length was found to increase with Mach number and decrease with interaction sweepback angle. For Mach numbers 3.0 to 5.0 the separation line was found to correspond to high intermittencies or equivalently to the downstream locus of the separation shock wave motion. The Mach 2.0 tests, however, showed that the intermittent region occurs significantly upstream of the separation line. Power spectral densities measured in the intermittent regions were found to have self-similar frequency distributions when compared as functions of a Strouhal number for all Mach numbers and interaction sweepback angles. The maximum zero-crossing frequencies were found to correspond with the peak frequencies in the power spectra measured in the intermittent region.

Barnhart, Paul J.↗
NASA NTRSDate not supplied

Transition Measurements on the SWiFT Model in the National Transonic Facility

Boundary layer transition locations on the Swept Wing Flow Test (SWiFT) Hybrid Wing Body model were measured in the National Transonic Facility (NTF) at the NASA Langley Research Center using several different methods. These methods include sublimating chemical flow visualization for a limited number of low Reynolds number conditions, as well as mean static pressure and dynamic pressure measurements. Results are presented for a range of conditions in the NTF, demonstrating Mach and Reynolds number effects. The transition front locations obtained from the different techniques agree well. A variation of the static pressure transition detection method is described in which continuous angle of attack polars are performed. This approach enables detection of the transition front at more challenging (high Reynolds number) conditions than is possible with sparser data from pitch-pause angle-of-attack polars. The trip dot effectiveness is also verified using all of the available diagnostics. Additionally, low Reynolds number results are compared with those acquired on the same model at the Aircraft Research Association (ARA) facility, and these results show good agreement overall, suggesting similar freestream flow quality conditions of the two facilities.

National Transonic Facility↗
NASA NTRSDate not supplied

Transition Measurements on the SWiFT Model in the National Transonic Facility

Boundary layer transition locations on the Swept Wing Flow Test (SWiFT) Hybrid Wing Body model were measured in the National Transonic Facility (NTF) at the NASA Langley Research Center using several different methods. These methods include sublimating chemical flow visualization for a limited number of low Reynolds number conditions, as well as mean static pressure and dynamic pressure measurements. Results are presented for a range of conditions in the NTF, demonstrating Mach and Reynolds number effects. The transition front locations obtained from the different techniques agree well. A variation of the static pressure transition detection method is described in which continuous angle of attack polars are performed. This approach enables detection of the transition front at more challenging (high Reynolds number) conditions than is possible with sparser data from pitch-pause angle-of-attack polars. The trip dot effectiveness is also verified using all of the available diagnostics. Additionally, low Reynolds number results are compared with those acquired on the same model at the Aircraft Research Association (ARA) facility, and these results show good agreement overall, suggesting similar freestream flow quality conditions of the two facilities.

National Transonic Facility↗
NASA NTRS1999

Tail Lobe Revisited: Magnetic Field Modeling Based on Plasma Data

Plasma data from the ISEE-1 and -2 spacecraft during 1977-1980 have been used to determine the distribution of data points in the magnetotail in the range of distances -20 < XGSM < --15, i.e. which of the records that were located in the current sheet, in the tail lobe, in the magnetosheath and in the boundary layers respectively. The ISEE-1 and -2 magnetic field data for the records in the tail lobe were then used to model the tail lobe magnetic field dependence on the solar wind dynamic pressure, on the Interplanetary Magnetic Field (IMF) and on the Dst index. The tail lobe magnetic field was assumed to be dependent on the square root of the dynamic pressure based on the balance between the total magnetic pressure in the tail lobes and the dynamic pressure of the solar wind. The IMF dependent terms, added to the pressure term, were sought in many different forms while the Dst dependence of the tail lobe magnetic field was assumed to be linear. The field shows a strong dependence on the square root of the dynamic pressure and the different IMF dependent terms all constitute a significant contribution to the total field. However, the dependence on the Dst index turned out to be very weak at those down-tail distances. The results of this study are intended to be used for parameterizing future versions of the data-based models of the global magnetospheric magnetic field.

Karlsson, S. B. P.↗
NASA NTRS1990

Entry vehicle performance analysis and atmospheric guidance algorithm for precision landing on Mars

Future missions to Mars may require pin-point landing precision, possibly on the order of tens of meters. The ability to reach a target while meeting a dynamic pressure constraint to ensure safe parachute deployment is complicated at Mars by low atmospheric density, high atmospheric uncertainty, and the desire to employ only bank angle control. The vehicle aerodynamic performance requirements and guidance necessary for 0.5 to 1.5 lift drag ratio vehicle to maximize the achievable footprint while meeting the constraints are examined. A parametric study of the various factors related to entry vehicle performance in the Mars environment is undertaken to develop general vehicle aerodynamic design requirements. The combination of low lift drag ratio and low atmospheric density at Mars result in a large phugoid motion involving the dynamic pressure which complicates trajectory control. Vehicle ballistic coefficient is demonstrated to be the predominant characteristic affecting final dynamic pressure. Additionally, a speed brake is shown to be ineffective at reducing the final dynamic pressure. An adaptive precision entry atmospheric guidance scheme is presented. The guidance uses a numeric predictor-corrector algorithm to control downrange, an azimuth controller to govern crossrange, and analytic control law to reduce the final dynamic pressure. Guidance performance is tested against a variety of dispersions, and the results from selected tests are presented. Precision entry using bank angle control only is demonstrated to be feasible at Mars.

Dieriam, Todd A.↗
NASA NTRS1988

Trajectory optimization and guidance law development for national aerospace plane applications

The work completed to date is comprised of the following: a simple vehicle model representative of the aerospace plane concept in the hypersonic flight regime, fuel-optimal climb profiles for the unconstrained and dynamic pressure constrained cases generated using a reduced order dynamic model, an analytic switching condition for transition to rocket powered flight as orbital velocity is approached, simple feedback guidance laws for both the unconstrained and dynamic pressure constrained cases derived via singular perturbation theory and a nonlinear transformation technique, and numerical simulation results for ascent to orbit in the dynamic pressure constrained case.

Calise, A. J.↗
NASA NTRS1990

Experimental transonic flutter characteristics of supersonic cruise configurations

The flutter characteristics of a generic arrow-wing supersonic transport configuration are studied. The wing configuration has a 3 percent biconvex airfoil and a leading-edge sweep of 73 deg out to a cranked tip with a 60 deg leading-edge sweep. The ground vibration tests and flutter test procedure are described. The effects of flutter on engine nacelles, fuel loading, wing-mounted vertical fin, wing angle-of-attack, and wing tip mass and stiffness distributions are analyzed. The data reveal that engine nacelles reduce the transonic flutter dynamic pressure by 25-30 percent; fuel loadings decrease dynamic pressures by 25 percent; 4-6 deg wing angles-of-attack cause steep transonic boundaries; and 5-10 percent changes in flutter dynamic pressures are the result of the wing-mounted vertical fin and wing-tip mass and stiffness distributions.

Durham, Michael H.↗
NASA NTRS2020

1- by 1-Foot Supersonic Wind Tunnel Sidewall Flowfield Compression Plate Study

A method for locally increasing the dynamic pressure in the test section of a supersonic wind tunnel is investigated from a theoretical, computational and experimental perspective, culminating in a reduced-scale, proof-of-concept test at Mach 3.44. Symmetric, variable-angle compression plates along the test section sidewalls generate oblique shockwaves to elevate the test section dynamic pressure (q). The experimental effort utilizes boundary-layer pitot rakes to sample the incoming flow near the floor, pitot and static rakes to measure the high-q core flow and numerous surface taps to map out the static pressure distribution along the floor, ceiling and sidewalls. A reliable and robust starting procedure is developed to establish supersonic flow in the test section and behind the compression plates. A characterization of the incoming floor boundary layer reveals previously documented features (such as a bulge toward the centerline) and an unexpected (and as-yet-unexplained) thickening of the boundary layer with increasing Reynolds number around total pressures of 40–60 psia. The high-q core flow is characterized for eight compression plate angles (4.0–8.0°), eight tunnel total pressures (20–139 psia) and three axial stations (24.95–27.35 inches). The peak dynamic pressure rise exceeds expectations. At conditions relevant to full-scale testing (6.0° and 139 psia), the dynamic pressure increases by a factor of 1.73, the Mach number becomes 2.85 and the total pressure drops imperceptibly. Basic distortion metrics show transverse non-uniformity of less than 5 %. Limited spatial resolution in the static rake measurements warrants additional, higher fidelity testing or computational analysis. Surface-mounted, fast-response instrumentation suggests minimal unsteadiness in the flow, less than a few percent of the mean (comparable to the sensor accuracy).

Wind tunnel testing, test techniques/methods, dyna↗
NASA NTRS1993

Six-degree-of-freedom guidance and control-entry analysis of the HL-20

The ability of the HL-20 lifting body to fly has been evaluated for an automated entry from atmospheric interface to landing. This evaluation was required to demonstrate that not only successful touchdown conditions would be possible for this low lift-to-drag-ratio vehicle, but also the vehicle would not exceed its design dynamic pressure limit of 400 psf during entry. This dynamic pressure constraint limit, coupled with limited available pitch-control authority at low supersonic speeds, restricts the available maneuvering capability for the HL-20 to acquire the runway. One result of this analysis was that this restrictive maneuvering capability does not allow the use of a model-following atmospheric entry-guidance algorithm, such as that used by the Space Shuttle, but instead requires a more adaptable guidance algorithm. Therefore, for this analysis, a predictor-corrector guidance algorithm was developed that would provide successful touchdown conditions while not violating the dynamic pressure constraint. A flight-control system was designed and incorporated, along with the predictor-corrector guidance algorithm, into a six-DOF simulation. which showed that the HL-20 remained controllable and could reach the landing site and execute a successful landing under all off-nominal conditions simulated.

Powell, Richard W.↗