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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 703 records · Page 39
All sourcesNASA NTRSDOE OSTI
NASA NTRS1958

Aerodynamic Characteristics in Sideslip of a Large-Scale 49 deg Sweptback Wing-Body-Tail Configuration with Blowing Applied Over the Flaps and Wing Leading Edge

An investigation has been conducted in the Langley full-scale tunnel to determine the aerodynamic characteristics in sideslip of a large-scale 490 sweptback wing-body-tail configuration having wing leading- edge and flap-blowing boundary-layer control. The wing and tails had an aspect ratio of 3.5, a taper ratio of 0.3, and NACA 65AO06 airfoil sections parallel to the plane of symmetry. The tests were conducted over a range of angles of attack of about -5 deg to 28 deg for sideslip angles of 0 deg, -5.06 deg, -10.15 deg, and -15.18 deg. Lateral and longitudinal stability and control characteristics were obtained for6a minimized blowing rate. The Reynolds number of the tests was 5.2 x 10(exp 6), corresponding to a Mach number of 0.08. The results of the investigation showed that sideslip to angles of about -15 deg did not require, from a consideration of the longitudinal characteristics, blowing rates over the wing leading edge or flap greater than that established as minimum at zero sideslip. The optimum configuration was laterally and directionally stable through the complete lift-coefficient range including the stall; however, maximum lift for sideslip angles greater than about 50 was seriously limited by a deficiency of lateral control. Blowing over the leading edge of the retreating wing in sideslip at a rate greater than that established as minimum at zero sideslip was ineffective in improving the lateral control characteristics. The optimum configuration at zero sideslip had no hysteresis of the aerodynamic parameters upon recovery from stall.

McLemore, H. Clyde↗
NASA NTRS1961

Static Lateral Characteristics at High Subsonic Speeds of a Complete Airplane Model with a Highly Tapered Wing having the 0.80 Chord Line Unswept and with Several Tail Configurations

An investigation was made at high subsonic speeds of a complete model having a highly tapered wing and several tail configurations. The aspect-ratio-3.50 wing had a taper ratio of 0.067 and an unswept 0.80 chord line. The complete model was tested with a wing-chord-plane tail, a T-tail, and a biplane tail (combined T-tail and wing-chord-plane tail). The model was tested in the Langley high-speed 7- by 10-foot tunnel at Mach numbers from 0.60 to 0.92 over a range of angle of attack of about +/- 20 deg. and a range of sideslip of -15 deg. to 13 deg. Some data were obtained with the horizontal stabilizer deflected. A few tests were also made with the wing tips clipped to an aspect ratio of 3.00. The data show that shock-interference effects between the tail surfaces (T-tail) can have considerable influence on the directional stability and effective dihedral. For example, the T-tail configuration with horizontal-tail leading-edge overhang showed a considerable loss in directional stability as the angle of attack was reduced to zero or negative values; whereas, the T-tail with zero leading-edge overhang showed the loss to be considerably less. The directional stability of the model with the low tail was essentially constant over a range of angle of attack of +/- 50 deg. All configurations tested showed a large reduction in stability at positive and negative angles of attack larger than about 15 deg., probably because of adverse sidewash associated with wing stall. The data show that a wing-chord-plane horizontal tail (low tail) tends to give a positive pitching-moment increment with increase in sideslip angle; whereas, a high tail (T-tail) tends to give negative increments in pitching moment.

Goodson, Kenneth W.↗
NASA NTRS1953

Drag Measurements on Equivalent Bodies of Revolution of Six Configurations of the Convair MX-1964 (Originally MX-1626) Proposed Supersonic Bomber

Tests on equivalent bodies of revolution of six configurations of the Consolidated Vultee Aircraft Corporation proposed supersonic bomber (Convair MX-1964) have indicated that it is possible to reduce the drag of the configuration by designing it to have a favorable area distribution. The method of NACA RM L53I22c to predict the peak pressure drag of a configuration on the basis of its area distribution gave generally good agreement with the subject models.

Hall, James Rudyard↗
NASA NTRS1998

Effect of Configuration Pitching Motion on Twin Tail Buffet Response

The effect of dynamic pitch-up motion of delta wing on twin-tail buffet response is investigated. The computational model consists of a delta wing-twin tail configuration. The computations are carried out on a dynamic multi-block grid structure. This multidisciplinary problem is solved using three sets of equations which consists of the unsteady Navier-Stokes equations, the aeroelastic equations, and the grid displacement equations. The configuration is pitched-up from zero up to 60 deg. angle of attack, and the freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. With the twin tail fixed as rigid surfaces and with no-forced pitch-up motion, the problem is solved for the initial flow conditions. Next, the problem is solved for the twin-tail response for uncoupled bending and torsional vibrations due to the unsteady loads on the twin tail and due to the forced pitch-up motion. The dynamic pitch-up problem is also solved for the flow response with the twin tail kept rigid. The configuration is investigated for inboard position of the twin tail which corresponds to a separation distance between the twin tail of 33% wing chord. The computed results are compared with the available experimental data.

Sheta, Essam F.↗
NASA NTRS1999

Ignition, Transition, Flame Spread in Multidimensional Configurations in Microgravity

In the inhabited quarters of orbiting spacecraft, fire is a greatly feared hazard. Thus, the fire safety strategy in a spacecraft is (1) to keep any fire as small as possible, (2) to detect any fire as early as possible, and (3) to extinguish any fire as quickly as possible. This suggests that a material which undergoes a momentary ignition might be tolerable but a material which permits a transition from a localized ignition to flame spread would significantly increase the fire hazard in a spacecraft. If the transition does not take place, fire growth does not occur. Therefore, it is critical to understand what process controls the transition. Many previous works have studied ignition and flame spread separately or were limited to a two-dimensional configuration. In this study, time-dependent phenomena of the transition over a thermally thin sample is studied experimentally and theoretically in two- and three-dimensional (2D,3D) configurations. Furthermore, localized ignition can be initiated at the center portion of thermally thin paper sample instead of at one end of the sample. Thus, the transition to flame spread could occur either toward upstream or downstream or both directions simultaneously with an external flow. In this presentation, the difference in the transition between the 3D and 2D configurations is explained with the numerically calculated data. For sufficiently narrow samples edge effects exist. Some results on this issue are presented. New analysis of the surface smoldering experiments conducted in the space shuttle STS-75 flight is also described.

Kashiwagi, Takashi↗
NASA NTRS1999

1999 NASA High-Speed Research Program Aerodynamic Performance Workshop: Configuration Aerodynamics - Volume 1

NASA's High-Speed Research Program sponsored the 1999 Aerodynamic Performance Technical Review on February 8-12, 1999 in Anaheim, California. The review was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in the areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High Lift, and Flight Controls. The review objectives were to: (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientists and engineers working on HSCT aerodynamics. In particular, single and midpoint optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT simulation results were presented, along with executive summaries for all the Aerodynamic Performance technology areas. The HSR Aerodynamic Performance Technical Review was held simultaneously with the annual review of the following airframe technology areas: Materials and Structures, Environmental Impact, Flight Deck, and Technology Integration. Thus, a fourth objective of the Review was to promote synergy between the Aerodynamic Performance technology area and the other technology areas of the HSR Program. This Volume 1/Part 1 publication covers configuration aerodynamics.

Hahne, David E.↗
NASA NTRS2000

Improving Jet Reactor Configuration for Production of Carbon Nanotubes

The jet mixing reactor has been proposed for the industrial production of fullerene carbon nanotubes. Here we study the flowfield of this reactor using the SIMPLER algorithm. Hot peripheral jets are used to enhance heating of the central jet by mixing with the ambiance of reactor. Numerous configurations of peripheral jets with various number of jets, distance between nozzles, angles between the central jet and a peripheral jets, and twisted configuration of nozzles are considered. Unlike the previous studies of jet mixing, the optimal configuration of peripheral jets produces strong non-uniformity of the central jet in a cross-section. The geometrical shape of reactor is designed to obtain a uniform temperature of a catalyst.

Povitsky, Alex↗
NASA NTRS2000

Fluid Mechanics, Drag Reduction and Advanced Configuration Aeronautics

This paper discusses Advanced Aircraft configurational approaches across the speed range, which are either enabled, or greatly enhanced, by clever Flow Control. Configurations considered include Channel Wings with circulation control for VTOL (but non-hovering) operation with high cruise speed, strut-braced CTOL transports with wingtip engines and extensive ('natural') laminar flow control, a midwing double fuselage CTOL approach utilizing several synergistic methods for drag-due-to-lift reduction, a supersonic strut-braced configuration with order of twice the L/D of current approaches and a very advanced, highly engine flow-path-integrated hypersonic cruise machine. This paper indicates both the promise of synergistic flow control approaches as enablers for 'Revolutions' in aircraft performance and fluid mechanic 'areas of ignorance' which impede their realization and provide 'target-rich' opportunities for Fluids Research.

Bushnell, Dennis M.↗
NASA NTRS2000

Development of an Integrated Nozzle for a Symmetric, RBCC Launch Vehicle Configuration

The development of rocket based combined cycle (RBCC) engines is highly dependent upon integrating several different modes of operation into a single system. One of the key components to develop acceptable performance levels through each mode of operation is the nozzle. It must be highly integrated to serve the expansion processes of both rocket and air-breathing modes without undue weight, drag, or complexity. The NASA GTX configuration requires a fixed geometry, altitude-compensating nozzle configuration. The initial configuration, used mainly to estimate weight and cooling requirements was a 1 So half-angle cone, which cuts a concave surface from a point within the flowpath to the vehicle trailing edge. Results of 3-D CFD calculations on this geometry are presented. To address the critical issues associated with integrated, fixed geometry, multimode nozzle development, the GTX team has initiated a series of tasks to evolve the nozzle design, and validate performance levels. An overview of these tasks is given. The first element is a design activity to develop tools for integration of efficient expansion surfaces With the existing flowpath and vehicle aft-body, and to develop a second-generation nozzle design. A preliminary result using a "streamline-tracing" technique is presented. As the nozzle design evolves, a combination of 3-D CFD analysis and experimental evaluation will be used to validate the design procedure and determine the installed performance for propulsion cycle modeling. The initial experimental effort will consist of cold-flow experiments designed to validate the general trends of the streamline-tracing methodology and anchor the CFD analysis. Experiments will also be conducted to simulate nozzle performance during each mode of operation. As the design matures, hot-fire tests will be conducted to refine performance estimates and anchor more sophisticated reacting-flow analysis.

Smith, Timothy D.↗
NASA NTRS2002

Investigation of Reynolds Number Effects on a Generic Fighter Configuration in the National Transonic Facility

A wind tunnel test of a generic fighter configuration was tested in the National Transonic Facility through a cooperative agreement between NASA Langley Research Center and McDonnell Douglas. The primary purpose of the test was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The test included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: (1) Fuselage/Wing; (2) Fuselage/Wing/Centerline Vertical Tail/Horizontal Tail; and (3) Fuselage/Wing/Trailing-Edge Extension/Twin Vertical Tails. Reynolds number effects on the longitudinal aerodynamic characteristics are presented herein.

Tomek, W. G.↗
NASA NTRS1996

Comparison of Aero/Propulsion Transition Characteristics for a Joint Strike Fighter Configuration

NASA Ames conducted a series tests of a large-scale powered model of Lockheed Martin Skunk Works' JSF STOVL configuration at the 80- by 120-Foot Wind Tunnel. The model uses a modified version of an F-16 fighter engine (a Pratt & Whitney F100-229 engine). To provide vertical flight capability, the Lockheed aircraft incorporates a forward fuselage mounted vertical lift fan in addition to an aft mounted vectoring nozzle all driven by a single engine. The lift fan is driven off the engine by means of a shaft and gear system. The data presented is for STOVL configuration 140, a wing-body-canard planform. The main goal of this testing was to obtain propulsion induced data for evaluating differences between small- and large-scale model testing during the transition flight regime between jet- and wing-borne flight. The data presented is from aerodynamic testing of the model in both the unpowered (engine-off) and powered (engine running) configurations. This data, when combined with the propulsion system thrust calibration data from the NASA Ames' Outdoor Aerodynamic Research Facility (OARF), is used to determine the propulsion induced effects during transition.

Naumowicz, Tim↗
NASA NTRS2001

Configuration Studies of Personal Air Vehicles. Personal Air Vehicle and Flying Jeep Concepts: A Commentary on Promising Approaches or What Goes Around Comes Around (About Every Twenty Years)

The NASA/Langley Personal Air Vehicle (PAV) Exploration (PAVE) and the DARPA (Defense Advanced Research Projects Agency) Dual Air/Road Transportation System (DARTS) projects were established to investigate the feasibility of creating vehicles which could replace, or at the very least augment, personal ground and air transportation schemes. This overall goal implies integrating several technology areas with practical everyday transportation requirements to design a class of vehicles which will achieve the following goals: (1) Vertical, Extremely Short, or Short Takeoff and Landing (VTOL, ESTOL, STOL) capability; (2) Operation at block speeds markedly faster than current combinations of land and air transportation, particularly in critical market areas; (3) Unit cost comparable to current luxury cars and small general aviation aircraft; (4) Excellent reliability; (5) Excellent safety; (6) Ability to integrate with existing land and air transportation systems. The conclusions of these configuration studies are summarized as follows: (1) Creation of the five assigned configurations prompted added explorations, some of which were dead-ends; (2) Some components could be common to all configurations such as avionics and dual-mode suspension schemes; (3) Single-Mode PAVs can be created by removing dual-mode-specific items; (4) Aviation history provided some intriguing starting points, as in what goes around comes around; (5) CTOL (Conventional Take-off and Landing) and STOL dual-mode PAVs look feasible with single-mode PAVs being simplifications of the dual-mode approach; (6) VTOL PAVs will require development; (7) More exotic collapsing mechanisms mechanisms need development; (8) As a teaching tool, PAVs are not yet a well-enough bounded design problem.

Hall, David W.↗
NASA NTRS2003

Investigation of Aerodynamics Scale Effects for a Generic Fighter Configuration in the National Transonic Facility (Invited)

Two wind tunnel tests of a generic fighter configuration have been completed in the National Transonic Facility. The primary purpose of the tests was to assess Reynolds number scale effects on a thin-wing, fighter-type configuration up to full-scale flight conditions (that is, Reynolds numbers of the order of 60 million). The tests included longitudinal and lateral/directional studies at subsonic and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to flight conditions. Results are presented for three Mach numbers (0.6, 0.8, and 0.9) and three configurations: 1) Fuselage / Wing, 2) Fuselage / Wing / Centerline Vertical Tail / Horizontal Tail, and 3) Fuselage / Wing / Trailing-Edge Extension / Twin Vertical Tails. Reynolds number effects on the lateral-directional aerodynamic characteristics are presented herein, along with longitudinal data demonstrating the effects of fixing the boundary layer transition location for low Reynolds number conditions. In addition, an improved model videogrammetry system and results are discussed.

Tomek, W. G.↗
NASA NTRS2003

Computational Aeroheating Predictions for Mars Lander Configurations

The proposed Mars Science Laboratory (MSL) mission is intended to deliver a large rover to the Martian surface within 10 km of the target site. This paper presents computational fluid dynamics (CFD) predictions of forebody heating rates for two MSL entry configurations with fixed aerodynamic trim tabs. Results are compared to heating on a 70-deg sphere-cone reference geometry. All three heatshield geometries are designed to trim hypersonically at a 16 deg angle of attack in order to generate the lift-to-drag ratio (LID) required for precision landing. Comparisons between CFD and tunnel data are generally in good agreement for each configuration, but the computations predict more flow separation and higher heating on a trim tab inclined 10 deg relative to the surface. CFD solutions at flight conditions were obtained using an 8-species Mars gas in chemical and thermal non-equilibrium. Laminar and Baldwin-Lomax solutions were used to estimate the effects of the trim tabs and turbulence on heating. A tab extending smoothly from the heatshield flank is not predicted to increase laminar or turbulent heating rates above the reference levels. Laminar heating on a tab deflected 10 deg from the conical heatshield is influenced by flow separation and is up to 35% above the baseline heating rate. The turbulent solution on the inclined tab configuration predicts attached flow and a 43% heating increase above the reference level.

Edquist, Karl T.↗
NASA NTRS2003

Computational Aeroheating Predictions for Mars Lander Configurations

The proposed Mars Science Laboratory (MSL) mission is intended to deliver a large rover to the Martian surface within 10 km of the target site. This paper presents computational fluid dynamics (CFD) predictions of forebody heating rates for two MSL entry configurations with fixed aerodynamic trim tabs. Results are compared to heating on a 70-deg sphere-cone reference geometry. All three heatshield geometries are designed to trim hypersonically at a 16 deg angle of attack in order to generate the lift-to-drag ratio (L/D) required for precision landing. Comparisons between CFD and tunnel data are generally in good agreement for each configuration, but the computations predict more flow separation and higher heating on a trim tab inclined 10 deg relative to the surface. CFD solutions at flight conditions were obtained using an 8-species Mars gas in chemical and thermal nonequilibrium. Laminar and Baldwin-Lomax solutions were used to estimate the effects of the trim tabs and turbulence on heating. A tab extending smoothly from the heatshield flank is not predicted to increase laminar or turbulent heating rates above the reference levels. Laminar heating on a tab deflected 10 deg from the conical heatshield is influenced by flow separation and is up to 35% above the baseline heating rate. The turbulent solution on the inclined tab configuration predicts attached flow and a 43% heating increase above the reference level.

Edquist, Karl T.↗
NASA NTRS2003

A Comparison of Active and Passive Methods for Control of Hypersonic Boundary Layers on Airbreathing Configurations

Active and passive methods for control of hypersonic boundary layers have been experimentally examined in NASA Langley Research Center wind tunnels on a Hyper-X model. Several configurations for forcing transition using passive discrete roughness elements and active mass addition, or blowing, methods were compared in two hypersonic facilities, the 20-Inch Mach 6 Air and the 31-Inch Mach 10 Air tunnels. Heat transfer distributions, obtained via phosphor thermography, shock system details, and surface streamline patterns were measured on a 0.333-scale model of the Hyper-X forebody. The comparisons between the active and passive methods for boundary layer control were conducted at test conditions that nearly match the nominal Mach 7 flight trajectory of an angle-of-attack of 2-deg and length Reynolds number of 5.6 million. For the passive roughness examination, the primary parametric variation was a range of trip heights within the calculated boundary layer thickness for several trip concepts. The prior passive roughness study resulted in a swept ramp configuration being selected for the Mach 7 flight vehicle that was scaled to be roughly 0.6 of the calculated boundary layer thickness. For the active jet blowing study, the blowing manifold pressure was systematically varied for each configuration, while monitoring the mass flow, to determine the jet penetration height with schlieren and transition movement with the phosphor system for comparison to the passive results. All the blowing concepts tested were adequate for providing transition onset near the trip location with manifold stagnation pressures on the order of 40 times the model static pressure or higher.

Berry, Scott A.↗
NASA NTRS2000

Analysis of Advanced Rotorcraft Configurations

Advanced rotorcraft configurations are being investigated with the objectives of identifying vehicles that are larger, quieter, and faster than current-generation rotorcraft. A large rotorcraft, carrying perhaps 150 passengers, could do much to alleviate airport capacity limitations, and a quiet rotorcraft is essential for community acceptance of the benefits of VTOL operations. A fast, long-range, long-endurance rotorcraft, notably the tilt-rotor configuration, will improve rotorcraft economics through productivity increases. A major part of the investigation of advanced rotorcraft configurations consists of conducting comprehensive analyses of vehicle behavior for the purpose of assessing vehicle potential and feasibility, as well as to establish the analytical models required to support the vehicle development. The analytical work of FY99 included applications to tilt-rotor aircraft. Tilt Rotor Aeroacoustic Model (TRAM) wind tunnel measurements are being compared with calculations performed by using the comprehensive analysis tool (Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics (CAMRAD 11)). The objective is to establish the wing and wake aerodynamic models that are required for tilt-rotor analysis and design. The TRAM test in the German-Dutch Wind Tunnel (DNW) produced extensive measurements. This is the first test to encompass air loads, performance, and structural load measurements on tilt rotors, as well as acoustic and flow visualization data. The correlation of measurements and calculations includes helicopter-mode operation (performance, air loads, and blade structural loads), hover (performance and air loads), and airplane-mode operation (performance).

Johnson, Wayne↗
NASA NTRS2004

An Inviscid Computational Study of the Space Shuttle Orbiter and Several Damaged Configurations

Inviscid aerodynamic characteristics of the Space Shuttle Orbiter were computed in support of the Columbia Accident Investigation. The unstructured grid software FELISA was used and computations were done using freestream conditions corresponding to those in the NASA Langley 20-Inch Mach 6 CF4 tunnel test section. The angle of attack was held constant at 40 degrees. The baseline (undamaged) configuration and a large number of damaged configurations of the Orbiter were studied. Most of the computations were done on a half model. However, one set of computations was done using the full-model to study the effect of sideslip. The differences in the aerodynamic coefficients for the damaged and the baseline configurations were computed. Simultaneously with the computation reported here, tests were being done on a scale model of the Orbiter in the 20-Inch Mach 6 CF4 tunnel to measure the deltas . The present computations complemented the CF4 tunnel test, and provided aerodynamic coefficients of the Orbiter as well as its components. Further, they also provided details of the flow field.

Prabhu, Ramadas K.↗