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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Development and testing of airfoils for high-altitude aircraft

Specific tasks included airfoil design; study of airfoil constraints on pullout maneuver; selection of tail airfoils; examination of wing twist; test section instrumentation and layout; and integrated airfoil/heat-exchanger tests. In the course of designing the airfoil, specifically for the APEX test vehicle, extensive studies were made over the Mach and Reynolds number ranges of interest. It is intended to be representative of airfoils required for lightweight aircraft operating at extreme altitudes, which is the primary research objective of the APEX program. Also considered were thickness, pitching moment, and off-design behavior. The maximum ceiling parameter M(exp 2)C(sub L) value achievable by the Apex-16 airfoil was found to be a strong constraint on the pullout maneuver. The NACA 1410 and 2410 airfoils (inverted) were identified as good candidates for the tail, with predictable behavior at low Reynolds numbers and good tolerance to flap deflections. With regards to wing twist, it was decided that a simple flat wing was a reasonable compromise. The test section instrumentation consisted of surface pressure taps, wake rakes, surface-mounted microphones, and skin-friction gauges. Also, a modest wind tunnel test was performed for an integrated airfoil/heat-exchanger configuration, which is currently on Aurora's 'Theseus' aircraft. Although not directly related to the APEX tests, the aerodynamics or heat exchangers has been identified as a crucial aspect of designing high-altitude aircraft and hence is relevant to the ERAST program.

Drela, Mark↗
NASA NTRS1976

Geared-elevator flutter study

An experimental and analytical study was made of the transonic flutter characteristics of a supersonic transport tail assembly model having an all-movable, horizontal tail with a geared elevator. Two model configurations, namely, one with a gear-elevator (2.8 to 1.0 gear ratio) and one with locked-elevator (1.0 to 1.0 gear ratio), were flutter tested in the Langley transonic dynamics tunnel with an empennage cantilever-mounted on a sting. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter dynamic pressure boundaries for both configurations were nearly flat over a Mach number range from 0.9 to 1.1. Flutter calculations (mathematical models) were made for the geared-elevator configuration using three subsonic lifting-surface methods. In one method, the elevator was treated as a discrete surface, and in the other two methods, the stabilizer and elevator were treated as a single warped-surface with the primary difference between these two methods being in the mathematical implementation used. A comparison of the experimental and analytical results shows that the discrete-elevator method predicted best the experimental flutter dynamic pressure level. However, the single warped-surface methods predicts more closely the experimental flutter frequencies and Mach number trends.

Ruhlin, C. L.↗
NASA NTRS2015

Oxide_Oxide Ceramic Matrix Composite (CMC) Exhaust Mixer Development in the NASA Environmentally Responsible Aviation (ERA) Project

Rolls-Royce North American Technologies, Inc. (LibertyWorksLW) began considering the development of CMC exhaust forced mixers in 2008, as a means of obtaining reduced weight and hotter operating temperature capability, while minimizing shape distortion during operation, which would improve mixing efficiency and reduce fuel burn. Increased component durability, enhanced ability to fabricate complex-shaped components, and engine noise reduction are other potential advantages of CMC mixers (compared to metallic mixers). In 2010, NASA was pursuing the reduction of NOx emissions, fuel burn, and noise from turbine engines in Phase I of the Environmentally Responsible Aviation (ERA) Project. ERA subtasks, including those focused on CMC components, were formulated with the goal of maturing technology from proof of concept validation (TRL 3) to a systemsubsystem or prototype demonstration in a relevant environment (TRL 6). In April 2010, the NASA Glenn Research Center (GRC) and LibertyWorks jointly initiated a CMC Exhaust System Validation Program within the ERA Project, teaming on CMC exhaust mixer development for subsonic jet engines capable of operating with increased performance. Our initial focus was on designing, fabricating, and characterizing the thrust and acoustic performance of a roughly quarter-scale 16-lobe oxide oxide CMC mixer and tail cone along with a conventional low bypass exhaust nozzle. Support Services, LLC (Allendale, MI) and ATK COI Ceramics, Inc. (COIC, in San Diego, CA) supported the design of a subscale nozzle assembly that consisted of an oxide oxide CMC mixer and center body, with each component mounted on a metallic attachment ring. That design was based upon the operating conditions a mixer would experience in a turbofan engine. Validation of the aerodynamic and acoustic performance of the subscale mixer via testing and the achievement of TRL 4 encouraged the NASALWCOIC team to move to the next phase where a full scale CMC mixer sized for a RR AE3007 engine and a compatible attachment flange were designed, followed by CMC component fabrication by COIC, and vibration testing at GRC under conditions simulating the structural and dynamic environment encountered during engine operation. AFRL (WPAFB) supported this testing by performing 3D laser vibrometry to identify the mixer mode shapes and modal frequencies. The successful fabrication and testing of such a component has been achieved. The CMC mixer demonstrated good durability during vibration testing at room and elevated temperature (TRL5). This has cleared the article for a ground-based test on a Rolls-Royce AE3007 engine, where the performance and benefits of the component can be further assessed.

exhaust nozzles↗
NASA NTRS1983

Solar powered aircraft

A cruciform wing structure for a solar powered aircraft is disclosed. Solar cells are mounted on horizontal wing surfaces. Wing surfaces with spanwise axis perpendicular to surfaces maintain these surfaces normal to the Sun's rays by allowing aircraft to be flown in a controlled pattern at a large bank angle. The solar airplane may be of conventional design with respect to fuselage, propeller and tail, or may be constructed around a core and driven by propeller mechanisms attached near the tips of the airfoils.

Phillips, W. H.↗
NASA NTRS2016

Innovative Flow Control Concepts for Drag Reduction

This paper highlights the technology development of two flow control concepts for aircraft drag reduction. The NASA Environmentally Responsible Aviation (ERA) project worked with Boeing to demonstrate these two concepts on a specially outfitted Boeing 757 ecoDemonstrator during the spring of 2015. The first flow control concept used Active Flow Control (AFC) to delay flow separation on a highly deflected rudder and increase the side force that it generates. This may enable a smaller vertical tail to provide the control authority needed in the event of an engine failure during takeoff and landing, while still operating in a conventional manner over the rest of the flight envelope. Thirty-one sweeping jet AFC actuators were installed and successfully flight-tested on the vertical tail of the 757 ecoDemonstrator. Pilot feedback, flow cone visualization, and analysis of the flight test data confirmed that the AFC is effective, as a smoother flight and enhanced rudder control authority were reported. The second flow control concept is the Insect Accretion Mitigation (IAM) innovation where surfaces were engineered to mitigate insect residue adhesion on a wing's leading edge. This is necessary because something as small as an insect residue on the leading edge of a laminar flow wing design can cause turbulent wedges that interrupt laminar flow, resulting in an increase in drag and fuel use. Several non-stick coatings were developed by NASA and applied to panels that were mounted on the leading edge of the wing of the 757 ecoDemonstrator. The performance of the coated surfaces was measured and validated by the reduction in the number of bug adhesions relative to uncoated control panels flown simultaneously. Both flow control concepts (i.e., sweeping jet actuators and non-stick coatings) for drag reduction were the culmination of several years of development, from wind tunnel tests to flight tests, and produced valuable data for the advancement of modern aircraft designs. The ERA systems analysis studies performed by NASA indicated that AFC-enhanced vertical tail could produce approximately 0.9% drag reduction for a large twin aisle aircraft and IAM coatings could enable approximately 1.2% drag reduction recovery for a potential total drag reduction of approximately 3.3% for a single aisle aircraft with a natural laminar flow (NLF) wing design.

Lin, John C.↗
NASA NTRSDate not supplied

Experimental Studies of Open Rotor Installation Effects

Open rotor propulsion technologies offer an opportunity for reducing fuel burn. At the same time, great challenges arise from this radically different engine architecture in terms of aircraft system integration. The aeroacoustic effects of integration are one of those key challenges and, furthermore, can offer the possibility for total aircraft system noise reduction. The inter-relationship of the aerodynamic and aeroacoustic system integration effects is particularly important to enable future application. An extensive model scale test campaign was conducted to investigate a broad range of these open rotor installation effects for both a conventional and an unconventional airframe. The experimental campaign was conducted in the Boeing Low Speed Aeroacoustic Facility with specially designed modifications for efficient positioning of the open rotor rig relative to the airframe. The airframe was traversed remotely relative to the fixed open rotor rig providing for the investigation of a large number of installation positions. Eight positions around the main wing of the conventional airframe and eleven positions above the hybrid wing body airframe were documented. Other parameters investigated in the test campaign included forward flight simulation Mach number, angle of attack, rotor speed, and airframe control surface deployment. In addition, both airframes were modified for alternate configurations. The conventional airframe was configured with both a T and a U tail while the unconventional hybrid wing body airframe was configured with different vertical control surfaces. An extensive instrumentation package was deployed. The acoustics were documented with a fixed array of far field microphones, a traversing array of in-flow microphones, an out-of-flow traversing phased array, and unsteady pressure transducers mounted flush on the surface of the airframes. In addition, mean flow surveys were measured with an articulating arm traversing system. The flow field surveys were particularly valuable in documenting flow distortion effects for the various installation positions and those created by angle of attack. This presentation will report the key results obtained for open rotor installation effects and discuss future prospects with the perspective of these results.

Michael J Czech↗
NASA NTRS1987

Long-range LFC transport

M = 0.83 Laminar Flow Control (LFC) transports, carrying large percentage payloads over a range of 20000 kilometers at cruise L/D's of 39 appear feasible with large space externally braced wings, external fuel pods, active controls, and 70 percent laminar flow on wing and tail surfaces, engine nacelles and struts, and a turbulent fuselage. A combination of a swept-forward inboard and a swept-back outer wing appears superior overall, especially for laminar flow and eliminating leading edge contamination probably caused by flyspecks and ice crystals. Wing divergence appears controllable by a combination of various methods. Wind-mounted superfans with extensive laminar flow on their nacelles appear practical. Their dominant tone noise is below the frequency range of the most strongly amplified TS-waves.

Pfenninger, Werner↗
NASA NTRS2000

Steady and Periodic Pressure Measurements on a Generic Helicopter Fuselage Model in the Presence of a Rotor

A wind tunnel test of a generic helicopter fuselage model with an independently mounted rotor has been conducted to obtain steady and periodic pressure data on the helicopter body. The model was tested at four advance ratios and three thrust coefficients. The periodic unsteady pressure coefficients are marked by four peaks associated with the passage of the four rotor blades. Blade passage effects are largest on the nose and tail boom of the model. The magnitude of the pulse increases with rotor thrust coefficient. Tabular listings of the unsteady pressure data are included to permit independent analysis. A CD-rom containing the steady and unsteady pressure data presented in the report is available from the authors.

Mineck, Raymond E.↗
NASA NTRS1999

Computational Assessment of Aft-Body Closure for the HSR Reference H Configuration

A study has been conducted to determine how well the USM3D unstructured Euler solver can be utilized to predict the flow over the High Speed Research (HSR) Reference H configuration with an ultimate goal of prediction of Sting interference so after body closure effects may be evaluated. This study has shown that the code can be used to predict the interference effects of a lower mounted blade sting with a high degree of confidence. It has been shown that wing and fuselage pressures, both levels and trends, can be predicted well. Force and moment levels are not predicted well but experimental trends are predicted. Based upon this, predicted force and moment increments are assumed to be predicted accurately. Deflection of the horizontal tail was found to cause a non-linear increment from the non-deflected sting interference effects.

Londenberg, W. Kelly↗
NASA NTRS1983

Effects of Nacelle configuration/position on performance of subsonic transport

An experimental study was conducted to explore possible reductions in installed propulsion system drag due to underwing aft nacelle locations. Both circular (C) and D inlet cross section nacelles were tested. The primary objectives were: to determine the relative installed drag of the C and D nacelle installations; and, to compare the drag of each aft nacelle installation with that of a conventional underwing forward, drag of each aft nacelle installation with that of a conventional underwing forward, pylon mounted (UTW) nacelle installation. The tests were performed in the NASA-Langley Research Center 16-Foot Transonic Wind Tunnel at Mach numbers from 0.70 to 0.85, airplane angles of attack from -2.5 to 4.1 degrees, and Reynolds numbers per foot from 3.4 to 4.0 million. The nacelles were installed on the NASA USB full span transonic transport model with horizontal tail on. The D nacelle installation had the smallest drag of those tested. The UTW nacelle installation had the largest drag, at 6.8 percent larger than the D at Mach number 0.80 and lift coefficient (C sub L) 0.45. Each tested configuration still had some interference drag, however. The effect of the aft nacelles on airplane lift was to increase C sub L at a fixed angle of attack relative to the wing body. There was higher lift on the inboard wing sections because of higher pressures on the wing lower surface. The effects of the UTW installation on lift were opposite to those of the aft nacelles.

Bangert, L. H.↗
NASA NTRS2011

Experimental Investigations of the NASA Common Research Model in the NASA Langley National Transonic Facility and NASA Ames 11-Ft Transonic Wind Tunnel (Invited)

Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility and the NASA Ames 11-ft wind tunnel. Data have been obtained at chord Reynolds numbers of 5 million for five different configurations at both wind tunnels. Force and moment, surface pressure and surface flow visualization data were obtained in both facilities but only the force and moment data are presented herein. Nacelle/pylon, tail effects and tunnel to tunnel variations have been assessed. The data from both wind tunnels show that an addition of a nacelle/pylon gave an increase in drag, decrease in lift and a less nose down pitching moment around the design lift condition of 0.5 and that the tail effects also follow the expected trends. Also, all of the data shown fall within the 2-sigma limits for repeatability. The tunnel to tunnel differences are negligible for lift and pitching moment, while the drag shows a difference of less than ten counts for all of the configurations. These differences in drag may be due to the variation in the sting mounting systems at the two tunnels.

Rivers, S. M.↗
NASA NTRS1947

Wind-Tunnel Tests of a 0.182-Scale Model of an F4U-1 Airplane with External Stores

Tests were made in the Langley 7- by 10-foot tunnel on a 0.182-scale model of an F4U-1 airplane with external stores. This paper is concerned mainly with presenting the data obtained in this investigation and with a comparison of some of these data with flight-test results determining the feasibility of estimating flight buffet Mach number from tunnel data. The results of this investigation indicate that the incremental drag coefficient due to external stores may be used to estimate the maximum Mach number that the F4U-1 airplane may reach in flight when it is equipped with external stores. This estimation is conservative for the five configurations investigated by mounts varying from 0 to 10 percent of the flight limit Mach number. The free-stream tunnel Mach number corresponding to sonic flow over the lower surface of the wing in the region of the store is a good indication of the lower limit of buffet in flight of the F4U-1 airplane when equipped with external stores. The fluctuations of total pressure over the horizontal tail are not sufficiently large (maximum of 1 percent q(sub o) to cause buffeting of the airplane.

Silvers, H. Norman↗
NASA NTRS1998

Aerothermal Effects of Cavities and Protuberances for High-Speed Sample Return Capsules

Extraterrestrial sample return is a growing component of solar system exploration. Currently, four missions, Stardust, 1 Muses-C, 2 Genesis, and Mars Sample Return, are under development that employ sample return as a prime component of the mission architecture. Respectively, these missions will return samples from the tail of a comet, an asteroid, the solar wind, and, Mars. An important component of these missions and the focus of this paper is the design of the sample return capsule (SRC). The purpose of the SRC is to safely return to Earth any gathered samples for terrestrial analysis. The two major design constraints for any SRC are as follows: 1) it must be able to survive a high-speed Earth entry (11 km/s to as a high as 15 km/s), 2) the mass of the SRC must be as small as possible. Because the SRC mass is carried from Earth to the sample sight and back, the SRC mass is a strong driver in the mission mass budget. Further, for the Mars Sample Return Capsule, planetary protection is another constraint. For this capsule, the probability of planetary contamination at Earth due to an SRC failure at entry must be minimal. For an SRC, a possible failure mechanism is severe local heating as a result of cavities and or protuberances in the SRC forebody heatshield. For example, the Apollo Command Module had a number of cavities and protuberances as part of the baseline designs Wind-tunnel tests of models containing small cavities and protuberances showed severe local heating augmentations in the vicinity of these surface discontinuities.4-5 As another example, the Genesis SRC forebody heat-shield contains penetrations (cavities) to mount the vehicle to the carrier bus. It is expected that these penetrations will also experience a severe local heating environment. A concern is that the large thermal gradients may produce sufficient thermal stress to cause local mechanical failure of the heatshield. Penetrations to the forebody heat-shield can also result from damage at vehicle integration, during launch, or during transportation of the sample return capsule from earth to the sample site and back. For example, the Starting SRC was damaged near the shoulder during the heatshield integration process producing a local surface discontinuity. Also, the Starting SRC traverses through the tail of a comet and is in space for 7 years. Thus, damage to the heatshield as a result of micrometeroid impact is a concern. Finally, it is difficult to characterize the effects of these potential heatshield singularities with ground-test facilities. Either detailed simulation or a dedicated flight test is required.

Olynick, David↗
NASA NTRS1975

Aerodynamic characteristics of a powered, externally blown flap STOL transport model with two engine simulator sizes

The low-speed aerodynamic characteristics are investigated of a general research model - a swept-wing, jet-powered STOL transport with externally blown flaps. The model was tested with four-engine simulators mounted on pylons under the 9.3-percent-thick supercritical airfoil wing. Two sets of air ejectors were used to provide data with large and small engines. Tests were conducted in the Langley V/STOL tunnel over an angle-of-attack range of -4 deg to 22 deg and a thrust-coefficient range from 0 to approximately 4. The effects are described of power, wing leading-edge slat configuration, T-tail and low horizontal-tail positions, and double-slotted flap deflection. Additional untrimmed and trimmed engine-out data and tail-body data are included.

Johnson, W. G., Jr.↗
NASA NTRS1977

Wind tunnel and ground static tests of a .094 scale powered model of a modified T-39 lift/cruise fan V/STOL research airplane

Ground static and wind tunnel test of a scale model modified T-39 airplane are reported. The configuration in the nose and replacement of the existing nacelles with tilting lift/cruise fans. The model was powered with three 14 cm diameter tip driven turbopowered simulators. Forces and moments were measured by an internal strain guage balance. Engine simulator thrust and mass flow were measured by calibrated pressure and temperature instrumentation mounted downstream of the fans. The low speed handling qualities and general aerodynamic characteristics of the modified T-39 were defined. Test variables include thrust level and thrust balance, forward speed, model pitch and sideslip angle at forward speeds, model pitch, roll, and ground height during static tests, lift/cruise fan tilt angle, flap and aileron deflection angle, and horizonal stabilizer angle. The effects of removing the landing gear, the lift/cruise fans, and the tail surfaces were also investigated.

Hunt, D.↗
NASA NTRS1976

Low-speed wind-tunnel investigation of a large scale advanced arrow-wing supersonic transport configuration with engines mounted above wing for upper-surface blowing

Tests have been conducted in a full scale tunnel to determine the low speed aerodynamic characteristics of a large scale advanced arrow wing supersonic transport configuration with engines mounted above the wing for upper surface blowing. Tests were made over an angle of attack range of -10 deg to 32 deg, sideslip angles of + or - 5 deg, and a Reynolds number range of 3,530,000 to 7,330,000. Configuration variables included trailing edge flap deflection, engine jet nozzle angle, engine thrust coefficient, engine out operation, and asymmetrical trailing edge boundary layer control for providing roll trim. Downwash measurements at the tail were obtained for different thrust coefficients, tail heights, and at two fuselage stations.

Shivers, J. P.↗
NASA NTRS1999

Supersonic Aftbody Closure Wind-Tunnel Testing, Data Analysis, and Computational Results

This paper reports on the model, test, and results from the Langley Supersonic Aftbody Closure wind tunnel test. This project is an experimental evaluation of the 1.5% Technology Concept Aircraft (TCA) aftbody closure model (Model 23) in the Langley Unitary Plan Wind Tunnel. The baseline TCA design is the result of a multidisciplinary, multipoint optimization process and was developed using linear design and analysis methods, supplemented with Euler and Navier-Stokes numerical methods. After a thorough design review, it was decided to use an upswept blade attached to the forebody as the mounting system. Structural concerns dictated that a wingtip support system would not be feasible. Only the aftbody part of the model is metric. The metric break was chosen to be at the fuselage station where prior aft-sting supported models had been truncated. Model 23 is thus a modified version of Model 20. The wing strongback, flap parts, and nacelles from Model 20 were used, whereas new aftbodies, a common forebody, and some new tails were fabricated. In summary, significant differences in longitudinal and direction stability and control characteristics between the ABF and ABB aftbody geometries were measured. Correcting the experimental data obtained for the TCA configuration with the flared aftbody to the representative of the baseline TCA closed aftbody will result in a significant reduction in longitudinal stability, a moderate reduction in stabilizer effectiveness and directional stability, and a moderate to significant reduction in rudder effectiveness. These reductions in the stability and control effectiveness levels of the baseline TCA closed aftbody are attributed to the reduction in carry-over area.

Allen, Jerry↗
NASA NTRS1946

Tests of a 1/5-Scale Model of the Republic XP-84 Airplane (Army Project MX-578) in the Langley 300 MPH 7- by 10-Foot Tunnel

A 1/5-scale model of the Republic x-84 airplane (Army Project MX-578) was tested in the Langley 300 MPH 7- by 10-foot tunnel. The primary object of the tests was twofold: to determine, a practicable method of increasing the longitudinal 3tability in the landing configuration, and to investigate the effects on longitudinal and lateral Stability of various external stored (fuel tanks, bombs, and rockets). The effects of the fuselage dive brakes were also determined, and the critical Mach numbers of certain of the airplane components were estimated. The use of the revised horizontal tail (of larger aspect ratio and area than the original) seemed to be the most feasible expedient for materially increasing the longitudinal stability in the landing configuration. The neutral-point shifts produced by the various external stores were unstable, the largest shift being about 2.5 percent mean aerodynamic chord. No appreciable aerodynamic trim changes were caused by the external stores. From the standpoint of range, maximum s peed, and rate of climb, the advantages of mounting the fuel tanks at the wing tips rather than inboard beneath the wings were clearly demonstrated by the tests. The effective dihedral parameter was the only static lateral-stability derivative appreciably affected by the external stores. At high lift coefficients, the tip-mounted tanks caused a large increase in the effective dihedral parameter (about 40 increase at a lift coefficient of 1.0). This increase was held undesirable, because the tendency toward oscillatory instability that it would cause would be heightened by the increased moments of inertia resulting from the weight of the tanks when carrying fuel. The fuselage dive brakes, when deflected, caused a change in trim tending to nose the airplane up; the neutral point also moved rearward upon deflecting the dive brakes. The amount of elevator required to overcome the change in trim was well within the available range of deflection. It was estimated that a drive-brake deflection of 900 would.decrease the terminal Mach number in a vertical dive by about 0.1. The estimated critical Mach number of the V-front canopy was about 0.04 greater than that of the original canopy. Pressure-distribution tests disclosed severe pressure peaks inside the nose of the jet entrance duct. These peaks, which would lead to separation and consequently poor pressure recovery at, the engine, could be reduced by, using a smaller nose,radius and: a modified internal lip shape

Tucker, Warren A.↗