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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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NASA NTRS2010

FASTSAT-HSV01 Synergistic Observations of the Magnetospheric Response During Active Periods: MINI-ME, PISA and TTI

Understanding the complex processes within the inner magnetosphere of Earth particularly during storm periods requires coordinated observations of the particle and field environment using both in-situ and remote sensing techniques. In fact in order to gain a better understanding of our Heliophysics and potentially improve our space weather forecasting capabilities, new observation mission approaches and new instrument technologies which can provide both cost effective and robust regular observations of magnetospheric activity and other space weather related phenomenon are necessary. As part of the effort to demonstrate new instrument techniques and achieve necessary coordinated observation missions, NASA's Fast Affordable Science and Technology Satellite Huntsville 01 mission (FASTSAT-HSVOI) scheduled for launch in 2010 will afford a highly synergistic solution which satisfies payload mission opportunities and launch requirements as well as contributing iri the near term to our improved understanding of Heliophysics. NASA's FASTSAT-HSV01 spacecraft on the DoD Space Test Program-S26 (STP-S26) Mission is a multi-payload mission executed by the DoD Space Test Program (STP) at the Space Development and Test Wing (SDTW), Kirtland AFB, NM. and is an example of a responsive and economical breakthrough in providing new possibilities for small space technology-driven and research missions. FASTSAT-HSV is a unique spacecraft platform that can carry multiple small instruments or experiments to low-Earth orbit on a wide range of expendable launch vehicles for a fraction of the cost traditionally required for such missions. The FASTSAT-HSV01 mission allows NASA to mature and transition a technical capability to industry while increasing low-cost access to space for small science and technology (ST) payloads. The FASTSAT-HSV01 payload includes three NASA Goddard Space Flight Center (GSFC) new technology built instruments that will study the terrestrial space environment and potentially contribute to space weather research in a synergistic manner. MINI-ME, a neutral atom imager, will observe the neutral atom inputs to ionospheric heating which can be important during high levels of magnetospheric activity. PISA, a plasma impedance spectrometer, will measure simultaneously the local electron densities and temperatures as well as measure small scale density structure (500 m spatial scale) during these active periods. TTI, a thermospheric imager, will remotely determine the thermospheric temperature response to this magnetospheric activity. Together, these observations will contribute significantly to a comprehensive understanding of the flow of energy through and the response of the storm-time terrestrial magnetosphere.

Casas, Joseph C.↗
NASA NTRS2010

Application of Approximate Unsteady Aerodynamics for Flutter Analysis

A technique for approximating the modal aerodynamic influence coefficient (AIC) matrices by using basis functions has been developed. A process for using the resulting approximated modal AIC matrix in aeroelastic analysis has also been developed. The method requires the unsteady aerodynamics in frequency domain, and this methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root locus et cetera. The unsteady aeroelastic analysis using unsteady subsonic aerodynamic approximation is demonstrated herein. The technique presented is shown to offer consistent flutter speed prediction on an aerostructures test wing (ATW) 2 and a hybrid wing body (HWB) type of vehicle configuration with negligible loss in precision. This method computes AICs that are functions of the changing parameters being studied and are generated within minutes of CPU time instead of hours. These results may have practical application in parametric flutter analyses as well as more efficient multidisciplinary design and optimization studies.

Pak, Chan-gi↗
NASA NTRS2010

Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

A technique for approximating the modal aerodynamic influence coefficients [AIC] matrices by using basis functions has been developed and validated. An application of the resulting approximated modal AIC matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO(TradeMark) flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing [ATW] 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle

Li, Wesley Waisang↗
NASA NTRS2011

Basis Function Approximation of Transonic Aerodynamic Influence Coefficient Matrix

A technique for approximating the modal aerodynamic influence coefficients matrices by using basis functions has been developed and validated. An application of the resulting approximated modal aerodynamic influence coefficients matrix for a flutter analysis in transonic speed regime has been demonstrated. This methodology can be applied to the unsteady subsonic, transonic, and supersonic aerodynamics. The method requires the unsteady aerodynamics in frequency-domain. The flutter solution can be found by the classic methods, such as rational function approximation, k, p-k, p, root-locus et cetera. The unsteady aeroelastic analysis for design optimization using unsteady transonic aerodynamic approximation is being demonstrated using the ZAERO flutter solver (ZONA Technology Incorporated, Scottsdale, Arizona). The technique presented has been shown to offer consistent flutter speed prediction on an aerostructures test wing 2 configuration with negligible loss in precision in transonic speed regime. These results may have practical significance in the analysis of aircraft aeroelastic calculation and could lead to a more efficient design optimization cycle.

Li, Wesley W.↗
NASA NTRS2020

Fiber-Optic Strain-Based Deflection and Twist Sensing for a High-Aspect-Ratio Swept Wing

Designs of aircraft structures have been moving toward leaner, lightweight designs for increased fuel efficiency. The Passive Aeroelastic Tailored (PAT) wing developed under the NASA Advanced Air Transport Technology (AATT) project is an example of a swept-wing design with high aspect ratio that incorporates lightweight highly-flexible tailored composite construction. The passive aeroelastic tailored structural design has explored the design space to enable aeroelastically tailored wing structures to increase aspect ratios (from 9 to 14) and ultimately reduce weight by 20 percent to 25 percent without impacting aeroelastic performance. To further study the aeroelastic performance of such a wing, the NASA Armstrong Flight Research Center (AFRC) (Edwards, California) has developed efficient real-time structural algorithms that are used in conjunction with a fiber-optic measurement system for lightweight vehicle applications. The AFRC Fiber Optic Sensing System (FOSS) provides up to 8,000 distributed surface strain measurements at one-half-inch increments and can be used to estimate a variety of structural parameters such as shape and load. This report discusses the implementation of strain-based displacement and twist-sensing techniques applied to the PAT wing test article tested at the NASA AFRC Flight Loads Laboratory. Empirical FOSS strain data are collected under varying loading conditions. Strain data are processed with the displacement and twist-sensing algorithms and independently verified by comparison to conventional ground-based instrumentation.

Pena, Francisco↗
NASA NTRS1982

Solar array Shuttle flight experiment - Hardware development and testing

This paper reports on the fabrication and ground testing of (a) a large area, light-weight, flexible substrate developmental solar array wing that has been built for NASA-MSFC and of (b) the supporting structure and data acquisition system (DAS) which, with the wing, will be flown in the Shuttle as an experiment in 1984. The experiment will verify the dynamics, thermodynamic, and electrical performance predictions of the array wing and will demonstrate the structural capability of the array wing for Orbiter launch and re-entry environments. The experiment hardware verification program was designed to minimize costs and risk of experiment performance degradation while maintaining Shuttle and crew safety. The previous full-scale wing hardware tests included an extension mast water table test and wing testing for random vibration, thermal vacuum, and acoustic environments. The results of these tests were used to define wing design modifications and to scope the test program for the experiment hardware.

Elms, R. V.↗
NASA NTRS1986

NASA rotor systems research aircraft: Fixed-wing configuration flight-test results

The fixed-wing, airplane configuration flight-test results of the Rotor System Research Aircraft (RSRA), NASA 740, at Ames/Dryden Flight Research Center are documented. Fourteen taxi and flight tests were performed from December 1983 to October 1984. This was the first time the RSRA was flown with the main rotor removed; the tail rotor was installed. These tests confirmed that the RSRA is operable as a fixed-wing aircraft. Data were obtained for various takeoff and landing distances, control sensitivity, trim and dynamics stability characteristics, performance rotor-hub drag, and acoustics signature. Stability data were obtained with the rotor hub both installed and removed. The speed envelope was developed to 261 knots true airspeed (KTAS), 226 knots calibrated airspeed (KCAS) at 10,000 ft density altitude. The airplane was configured at 5 deg. wing incidence with 5 deg. wing flaps as a normal configuration. Level-flight data were acquired at 167 KCAS for wing incidence from 0 to 10 deg. Step inputs and doublet inputs of various magnitudes were utilized to acquire dynamic stability and control sensitivity data. Sine-wave inputs of constantly increasing frequency were used to generate parameter identification data. The maximum load factor attained was 2.34 g at 206 KCAS.

Erickson, R. E.↗
NASA NTRS1977

Prediction of in-depth gap heating ratios from wing glove model test data

In-depth gap heating ratios were predicted down RSI tile sidewalls based on temperature measurements obtained from the JSC arc-jet Wing Glove model tests in order to develop gap heating ratios which resulted in the best possible fit of test data and to produce a set of engineering verification heating ratios similar in shape to one another which could be used at various body points on the Orbiter during reentry. The Rockwell TPS Multidimensional heat conduction program was used to perform 3-D thermal analyses using a 3.0 in. thick section of a curved RSI tile with 283 nodal points. Correlation with test data shows that the predicted heating ratios were significantly higher down in the gap than the zero pressure values for T/C stacks 39 and 38 on the Wing Glove model. For stack 37 (in a low pressure region), the baseline heating ratio overpredicted the temperature data. This analysis, which showed that the heating ratios were a strong function of the product of pressure and pressure gradient, will be used to compare with recent Gap/Step and Ames Double Wedge test/analysis results in the effort to identify the Orbiter gap response to high delta P flight environment.

Source record↗
NASA NTRS1959

Low-Speed Tests of Semispan-Wing Models at Angles of Attack from 0 to 180 degrees

Semispan-wing models were tested at angles of attack from 0 to 180 deg at low subsonic speeds. Eight plan forms were considered, both swept and unswept with aspect ratios ranging from 2 to 6. Except for a delta-wing model of aspect ratio 2. all models had a taper ratio of 0.5 and an NACA 64AO10 airfoil section. The delta-wing model had an NACA 0005 (modified) airfoil section. With two exceptions, the models were tested both with and without a full-span trailing-edge flap deflected 25 deg. The Reynolds numbers based on the mean aerodynamic chord were between 1.5 and 2.2 million. Lift, drag, and pitching-moment coefficients are presented as functions of angle of attack. Approximate corrections for the effects of blockage were applied to the data.

Koenig, David G.↗
NASA NTRS1960

Transonic and Supersonic Wind-Tunnel Tests of Wing-Body Combinations Designed for High Efficiency at a Mach Number of 1.41

Wind-tunnel force tests of a number of wing-body combinations designed for high lift-drag ratio at a Mach number of 1.41 are reported. Five wings and six bodies were used in making up the various wing-body combinations investigated. All the wings had the same highly swept dis- continuously tapered plan form with NACA 65A-series airfoil sections 4 percent thick at the root tapering linearly to 3 percent thick at the tip. The bodies were based on the area distribution of a Sears-Haack body of revolution for minimum drag with a given length and volume. These wings and bodies were used to determine the effects of wing twist., wing twist and camber, wing leading-edge droop, a change from circular to elliptical body cross-sectional shape, and body indentation by the area-rule and streamline methods. The supersonic test Mach numbers were 1.41 and 2.01. The transonic test Mach number range was from 0.6 to 1.2. For the transition-fixed condition and at a Reynolds number of 2.7 x 10(exp 6) based on the mean aerodynamic chord, the maximum value of lift- drag ratio at a Mach number of 1.41 was 9.6 for a combination with a twisted wing and an indented body of elliptical cross section. The tests indicated that the transonic rise in minimum drag was low and did not change appreciably up to the highest test Mach number of 2.01. The lower values of lift-drag ratio obtained at a Mach number of 2.01 can be attributed to the increase of drag due to lift with Mach number.

Grant, Frederick C.↗
NASA NTRS2018

Spanwise Adaptive Wing - PTERA Flight Test

This is a technology which depends on a morphing vehicle wingtip. It allows one to replace the use of the rudder for some vehicles with the outer ailerons for control of the Dutch Roll Mode. Description of the PTERA flight test in support of the Spanwise Adaptive Wing flight research program.

wing design↗
NASA NTRS1974

Static noise tests on augmentor wing jet STOL research aircraft (C8A Buffalo)

Results are presented for full scale ground static acoustic tests of over-area conical nozzles and a lobe nozzle installed on the Augmentor Wing Jet STOL Research Aircraft, a modified C8A Buffalo. The noise levels and spectrums of the test nozzles are compared against those of the standard conical nozzle now in use on the aircraft. Acoustic evaluations at 152 m (500 ft), 304 m (1000 ft), and 1216 m (4000 ft) are made at various engine power settings with the emphasis on approach and takeoff power. Appendix A contains the test log and propulsion calculations. Appendix B gives the original test plan, which was closely adhered to during the test. Appendix C describes the acoustic data recording and reduction systems, with calibration details.

Marrs, C. C.↗
NASA NTRS1934

Wing-Nacelle-Propeller Tests - Comparative Tests of Liquid-Cooled and Air-Cooled Engine Nacelles

This report gives the results of measurements of the lift, drag, and propeller characteristics of several wing and nacelle combinations with a tractor propeller. The nacelles were so located that the propeller was about 31% of the wing chord directly ahead of the leading edge of the wing, a position which earlier tests (NASA Report No. 415) had shown to be efficient. The nacelles were scale models of an NACA cowled nacelle for a radial air-cooled engine, a circular nacelle with the V-type engine located inside and the radiator for the cooling liquid located inside and the radiator for the type, and a nacelle shape simulating the housing which would be used for an extension shaft if the engine were located entirely within the wing. The propeller used in all cases was a 4-foot model of Navy No. 4412 adjustable metal propeller. The results of the tests indicate that, at the angles of attack corresponding to high speeds of flight, there is no marked advantage of one type of nacelle over the others as far as low drag is concerned, since the drag added by any of the nacelles in the particular location ahead of the wing is very small. The completely cowled nacelle for a radial air-cooled engine appears to have the highest drag, the liquid-cooled engine appears to have the highest drag, the liquid-cooled engine nacelle with external radiator slightly less drag. The liquid-cooled engine nacelle with radiator in the cowling hood has about half the drag of the cowled radial air-cooled engine nacelle. The extension-shaft housing shows practically no increase in drag over that of the wing alone. A large part of the drag of the liquid-cooled engine nacelle appears to be due to the external radiator. The maximum propulsive efficiency for a given propeller pitch setting is about 2% higher for the liquid-cooled engine nacelle with the radiator in the cowling hood than that for the other cowling arrangements.

Wood, Donald H.↗
NASA NTRS2017

Aeroelastic Modeling of X-56A Stiff-Wing Configuration Flight Test Data

Aeroelastic stability and control derivatives for the X-56A Multi-Utility Technology Testbed (MUTT), in the stiff-wing configuration, were estimated from flight test data using the output-error method. Practical aspects of the analysis are discussed. The orthogonal phase-optimized multisine inputs provided excellent data information for aeroelastic modeling. Consistent parameter estimates were determined using output error in both the frequency and time domains. The frequency domain analysis converged faster and was less sensitive to starting values for the model parameters, which was useful for determining the aeroelastic model structure and obtaining starting values for the time domain analysis. Including a modal description of the structure from a finite element model reduced the complexity of the estimation problem and improved the modeling results. Effects of reducing the model order on the short period stability and control derivatives were investigated.

Grauer, Jared A.↗
NASA NTRS1972

EBF noise tests with engine under-the-wing and over-the-wing configurations

Noise tests of externally blown flaps with the engine under the wing and engine over the wing configurations were conducted. Flap noise data obtained on a TF-34 aircraft are discussed. Noise data obtained during a free-jet forward-speed-effect analysis are presented. Noise sources associated with upper surface flap blowing are described. Results of a small scale configuration screening study and some large scale model test data are analyzed. The noise data for the engine over wing configurations are compared with the engine under the wing configurations.

Dorsch, R. G.↗
NASA NTRS1955

Full-scale-wind-tunnel Tests of a 35 Degree Sweptback Wing Airplane with High-velocity Blowing over the Training-edge Flaps

A wind-tunnel investigation was made to determine the effects of ejecting high-velocity air near the leading edge of plain trailing-edge flaps on a 35 degree sweptback wing. The tests were made with flap deflections from 45 degrees to 85 degrees and with pressure ratios across the flap nozzles from sub-critical up to 2.9. A limited study of the effects of nozzle location and configuration on the efficiency of the flap was made. Measurements of the lift, drag, and pitching moment were made for Reynolds numbers from 5.8 to 10.1x10(6). Measurements were also made of the weight rate of flow, pressure, and temperature of the air supplied to the flap nozzles.The results show that blowing on the deflected flap produced large flap lift increments. The amount of air required to prevent flow separation on the flap was significantly less than that estimated from published two-dimensional data. When the amount of air ejected over the flap was just sufficient to prevent flow separation, the lift increment obtained agreed well with linear inviscid fluid theory up to flap deflections of 60 degrees. The flap lift increment at 85 degrees flap deflection was about 80 percent of that predicted theoretically.With larger amounts of air blown over the flap, these lift increments could be significantly increased. It was found that the performance of the flap was relatively insensitive to the location of the flap nozzle, to spacers in the nozzle, and to flow disturbances such as those caused by leading-edge slats or discontinuities on the wing or flap surfaces. Analysis of the results indicated that installation of this system on an F-86 airplane is feasible.

Kelley, Mark W↗