Space shuttle: Heat transfer rate distributions on McDonnell-Douglas delta wing orbiter determined by phase-change paint technique for nominal Mach number of 8

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

The results are reported of the phase-change paint tests conducted at Mach 8, to determine the aerodynamic heat transfer distributions on the McDonnell Douglas delta wing orbiter. Model details, test conditions, and reduced heat transfer data are presented.

Unsteady, Transonic Flow Around Delta Wings Undergoing Coupled and Natural Modes Response: A Multidisciplinary Problem

NASA Technical Reports Server (NTRS)

Menzies, Margaret Anne

1996-01-01

The unsteady, three-dimensional Navier-Stokes equations coupled with the Euler equations of rigid-body dynamics are sequentially solved to simulate and analyze the aerodynamic response of a high angle of attack delta wing undergoing oscillatory motion. The governing equations of fluid flow and dynamics of the multidisciplinary problem are solved using a time-accurate solution of the laminar, unsteady, compressible, full Navier- Stokes equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The primary model under consideration consists of a 65 deg swept, sharp-edged, cropped delta wing of zero thickness at 20 deg angle of attack. In a freestream of Mach 0.85 and Reynolds number of 3.23 x 10(exp 6), the flow over the upper surface of the wing develops a complex shock system which interacts with the leading-edge primary vortices producing vortex breakdown. The effect of the oscillatory motion of the wing on the vortex breakdown and overall aerodynamic response is detailed to provide insight to the complicated physics associated with unsteady flows and the phenomenon of wing rock. Forced sinusoidal single and coupled mode rolling and pitching motion is presented for the wing in a transonic freestream. The Reynolds number, frequency of oscillation, and the phase angle are varied. Comparison between the single and coupled mode forced rolling and pitching oscillation cases illustrate the effects of coupling the motion. This investigation shows that even when coupled, forced rolling oscillation at a reduced frequency of 2(pi) eliminates the vortex breakdown which results in an increase in lift. The coupling effect for in phase forced oscillations show that the lift coefficient of the pitching-alone case and the rolling-moment coefficient of the rolling-alone case dominate the resulting response. However, with a phase lead in the pitching motion, the coupled motion results in a non-periodic response of the rolling moment. The second class of problems involve releasing the wing in roll to respond to the flowfield. Two models of sharp-edged delta wings, the previous 65 deg swept model and an 80 deg swept, sharp-edged delta wing, are used to observe the aerodynamic response of a wing free to roll in a transonic and subsonic freestream, respectively. These cases demonstrate damped oscillations, self-sustained limit cycle oscillations, and divergent rolling oscillations. Ultimately, an active control model using a mass injection system was applied on the surface of the wing to suppress the self-sustained limit cycle oscillation known as wing rock. Comparisons with experimental investigations complete this study, validating the analysis and illustrating the complex details afforded by computational investigations.

A Study of Grid Resolution, Transition and Turbulence Model Using the Transonic Simple Straked Delta Wing

NASA Technical Reports Server (NTRS)

Bartels, Robert E.

2001-01-01

Three-dimensional transonic flow over a delta wing is investigated using several turbulence models. The performance of linear eddy viscosity models and an explicit algebraic stress model is assessed at the start of vortex flow, and the results compared with experimental data. To assess the effect of transition location, computations that either fix transition aft of the leading edge or are fully turbulent are performed. These computations show that grid resolution, transition location and turbulence model significantly affect the 3D flowfield.

Hypersonic shock wave interaction and impingement

NASA Technical Reports Server (NTRS)

Kessler, W. C.; Reilly, J. F.; Sampatacos, E.

1971-01-01

An experimental investigation was conducted on space shuttle type, body-wing configurations. The purpose of the investigation was to determine the effects of body and wing geometry on the hypersonic shock structure about these vehicles and on the resulting surface impingement of interior flow field shock and expansion waves. Schlieren photographs and thermographic phosphor paint data were obtained on three body cross sections with three wing planforms at 40, 50 and 60 degree angles of attack. Specific configuration data were obtained at 0 and 30 degree angles of attack to develop trends. These data were obtained at a nominal Mach number of 13.5 and a freestream unit Reynolds number of 0.7 million per foot. For comparison with these straight wing configurations, data were also obtained on a model of a point design, high cross-range, delta wing orbiter at 40, 50 and 60 degree angles of attack. As expected, the data on this delta wing orbiter indicated that the shock intersection/impingement phenomena associated with straight wing vehicles are considerably more complex than, and result in both windward and leeward surface heating regions not present on, the delta configuration.

CFD validation experiments at McDonnell Aircraft Company

NASA Technical Reports Server (NTRS)

Verhoff, August

1987-01-01

Information is given in viewgraph form on computational fluid dynamics (CFD) validation experiments at McDonnell Aircraft Company. Topics covered include a high speed research model, a supersonic persistence fighter model, a generic fighter wing model, surface grids, force and moment predictions, surface pressure predictions, forebody models with 65 degree clipped delta wings, and the low aspect ratio wing/body experiment.

Three-dimensional aerodynamic shape optimization of supersonic delta wings

NASA Technical Reports Server (NTRS)

Burgreen, Greg W.; Baysal, Oktay

1994-01-01

A recently developed three-dimensional aerodynamic shape optimization procedure AeSOP(sub 3D) is described. This procedure incorporates some of the most promising concepts from the area of computational aerodynamic analysis and design, specifically, discrete sensitivity analysis, a fully implicit 3D Computational Fluid Dynamics (CFD) methodology, and 3D Bezier-Bernstein surface parameterizations. The new procedure is demonstrated in the preliminary design of supersonic delta wings. Starting from a symmetric clipped delta wing geometry, a Mach 1.62 asymmetric delta wing and two Mach 1. 5 cranked delta wings were designed subject to various aerodynamic and geometric constraints.

Experimental investigation of high-incidence delta-wing flow control

NASA Astrophysics Data System (ADS)

Buzica, Andrei; Bartasevicius, Julius; Breitsamter, Christian

2017-09-01

The possibility of extending the flight envelope for configurations with slender delta-shaped wings is investigated in this study by means of active flow control through pulsating jets from slot pairs distributed along the leading edge. The experiments comprise stereoscopic particle image velocimetry as well as force and moment measurements on a half-delta wing model. The analysis focuses on three high-incidence regimes: pre-stall, stall, and post-stall. This study also compares different perturbation methods: blowing with spatially constant and variable parameters, frequency and phase. At an incidence of 45°, the unison pulsed blowing facilitates the most significant flow transformation. Here, the separated shear layer reattaches on the wing's suction side, thus increasing the lift. Phase-averaged flow field measurements describe, in this particular case, the underlying physics of the flow-disturbance interaction.

Space shuttle phase B wind tunnel model and test information. Volume 2: Orbiter configuration

NASA Technical Reports Server (NTRS)

Glynn, J. L.; Poucher, D. E.

1988-01-01

Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a data base and are available for applying to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Data Base is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retro-glide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configuration types include booster and orbiter components in various stacked and tandem combinations.

Space shuttle phase B wind tunnel model and test information. Volume 3: Launch configuration

NASA Technical Reports Server (NTRS)

Glynn, J. L.; Poucher, D. E.

1988-01-01

Archived wind tunnel data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel data acquired in the Phase B development have been compiled into a data base and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include booster, orbiter and launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbital configuration types include straight and delta wings, lifting body, drop tanks and double delta wings. This is Volume 3 (Part 2) of the report -- Launch Configuration -- which includes booster and orbiter components in various stacked and tandem combinations.

Heat transfer investigation of two Langley Research Center delta wing configurations at a Mach number of 10.5, volume 1

NASA Technical Reports Server (NTRS)

Eaves, R. H.; Buchanan, T. D.; Warmbrod, J. D.; Johnson, C. B.

1972-01-01

Heat transfer tests for two delta wing configurations were conducted in the hypervelocity wind tunnel. The 24-inch long models were tested at a Mach number of approximately 10.5 and at angles of attack of 20, 40, and 60 degrees over a length Reynolds number range from 5 million to 23 million on 4 May to 4 June 1971. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint.

Effective Control of Computationally Simulated Wing Rock in Subsonic Flow

NASA Technical Reports Server (NTRS)

Kandil, Osama A.; Menzies, Margaret A.

1997-01-01

The unsteady compressible, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the delta wing rock phenomenon. The NS equations are solved time accurately, using the implicit, upwind, Roe flux-difference splitting, finite-volume scheme. The rigid-body dynamics equations are solved using a four-stage Runge-Kutta scheme. Once the wing reaches the limit-cycle response, an active control model using a mass injection system is applied from the wing surface to suppress the limit-cycle oscillation. The active control model is based on state feedback and the control law is established using pole placement techniques. The control law is based on the feedback of two states: the roll-angle and roll velocity. The primary model of the computational applications consists of a 80 deg swept, sharp edged, delta wing at 30 deg angle of attack in a freestream of Mach number 0.1 and Reynolds number of 0.4 x 10(exp 6). With a limit-cycle roll amplitude of 41.1 deg, the control model is applied, and the results show that within one and one half cycles of oscillation, the wing roll amplitude and velocity are brought to zero.

Space shuttle phase B wind tunnel model and test information. Volume 1: Booster configuration

NASA Technical Reports Server (NTRS)

Glynn, J. L.; Poucher, D. E.

1988-01-01

Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter, and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. This is Volume 1 (Part 2) of the report -- Booster Configuration.

Space shuttle phase B wind tunnel model and test information. Volume 1: Booster configuration

NASA Technical Reports Server (NTRS)

Glynn, J. L.; Poucher, D. E.

1988-01-01

Archived wind tunnel test data are available for flyback booster or other alternative recoverable configurations as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle. Considerable wind tunnel data was acquired by the competing contractors and the NASA Centers for an extensive variety of configurations with an array of wing and body planforms. All contractor and NASA wind tunnel test data acquired in the Phase B development have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration type. Basic components include the booster, the orbiter and the launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configuration types include straight and delta wings, lifting body, drop tanks, and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. This is Volume 1 (Part 1) of the report -- Booster Configuration.

Space shuttle phase B wind tunnel model and test information. Volume 3: Launch configuration

NASA Technical Reports Server (NTRS)

Glynn, J. L.; Poucher, D. E.

1988-01-01

Archived wind tunnel test data are available for flyback booster or other alternate recoverable configuration as well as reusable orbiters studied during initial development (Phase B) of the Space Shuttle, including contractor data for an extensive variety of configurations with an array of wing and body planforms. The test data have been compiled into a database and are available for application to current winged flyback or recoverable booster aerodynamic studies. The Space Shuttle Phase B Wind Tunnel Database is structured by vehicle component and configuration. Basic components include booster, orbiter, and launch vehicle. Booster configuration types include straight and delta wings, canard, cylindrical, retroglide and twin body. Orbiter configurations include straight and delta wings, lifting body, drop tanks and double delta wings. Launch configurations include booster and orbiter components in various stacked and tandem combinations. The digital database consists of 220 files containing basic tunnel data. Database structure is documented in a series of reports which include configuration sketches for the various planforms tested. This is Volume 3 -- launch configurations.

Supersonic aerodynamics of delta wings

NASA Technical Reports Server (NTRS)

Wood, Richard M.

1988-01-01

Through the empirical correlation of experimental data and theoretical analysis, a set of graphs has been developed which summarize the inviscid aerodynamics of delta wings at supersonic speeds. The various graphs which detail the aerodynamic performance of delta wings at both zero-lift and lifting conditions were then employed to define a preliminary wing design approach in which both the low-lift and high-lift design criteria were combined to define a feasible design space.

Effect of spanwise blowing on leading-edge vortex bursting of a highly swept aspect ratio 1.18 delta wing

NASA Technical Reports Server (NTRS)

Scantling, W. L.; Gloss, B. B.

1974-01-01

An investigation was conducted in the Langley 1/8-scale V/STOL model tunnel on a semispan delta wing with a leading-edge sweep of 74 deg, to determine the effectiveness of various locations of upper surface and reflection plane blowing on leading-edge vortex bursting. Constant area nozzles were located on the wing upper surface along a ray swept 79 deg, which was beneath the leading-edge vortex core. The bursting and reformation of the leading-edge vortex was viewed by injecting helium into the vortex core, and employing a schlieren system.

Fundamental aerodynamic characteristics of delta wings with leading-edge vortex flows

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.

1985-01-01

An investigation of the aerodynamics of sharp leading-edge delta wings at supersonic speeds has been conducted. The supporting experimental data for this investigation were taken from published force, pressure, and flow-visualization data in which the Mach number normal to the wing leading edge is always less than 1.0. The individual upper- and lower-surface nonlinear characteristics for uncambered delta wings are determined and presented in three charts. The upper-surface data show that both the normal-force coefficient and minimum pressure coefficient increase nonlinearly with a decreasing slope with increasing angle of attack. The lower-surface normal-force coefficient was shown to be independent of Mach number and to increase nonlinearly, with an increasing slope, with increasing angle of attack. These charts are then used to define a wing-design space for sharp leading-edge delta wings.

Skin friction fields on delta wings

NASA Astrophysics Data System (ADS)

Woodiga, S. A.; Liu, Tianshu

2009-12-01

The normalized skin friction fields on a 65° delta wing and a 76°/40° double-delta wing are measured by using a global luminescent oil-film skin friction meter. The detailed topological structures of skin friction fields on the wings are revealed for different angles of attack and the important features are detected such as reattachment lines, secondary separation lines, vortex bursting and vortex interaction. The comparisons with the existing flow visualization results are discussed.

77 FR 24835 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-04-26

... inspections must be used for that wing for that visit and for all subsequent repeat inspections. Delta stated... Company Model 767 airplanes. This AD was prompted by reports of cracking in the upper wing skin at the... in the loss of the strut-to-wing upper link load path and possible separation of a strut and engine...

Space shuttle: Heat transfer investigation of the McDonnell-Douglas delta wing orbiter at a nominal Mach number of 10.5

NASA Technical Reports Server (NTRS)

Eaves, R. H.; Buchanan, T. D.

1972-01-01

Heat transfer tests for the delta wing orbiter were conducted in a hypervelocity wind tunnel. A 1.1 percent scale model was tested at a Mach number of approximately 10.5 over an angle of attack range from 10 to 60 degrees over a length Reynolds number range from 5 times 10 to the 6th power to 24 times 10 to the 6th power. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint. Limited pressure measurements were obtained.

Visualization of leading edge vortices on a series of flat plate delta wings

NASA Technical Reports Server (NTRS)

Payne, Francis M.; Ng, T. Terry; Nelson, Robert C.

1991-01-01

A summary of flow visualization data obtained as part of NASA Grant NAG2-258 is presented. During the course of this study, many still and high speed motion pictures were taken of the leading edge vortices on a series of flat plate delta wings at varying angles of attack. The purpose is to present a systematic collection of photographs showing the state of vortices as a function of the angle of attack for the four models tested.

A computational study of incipient leading-edge separation on a 65-deg delta wing at M = 1.60

NASA Technical Reports Server (NTRS)

Mcmillin, S. Naomi; Pittman, James L.; Thomas, James L.

1990-01-01

A computational study on a 65-deg delta wing at a freestream Mach number of 1.60 has been conducted by obtaining conical Reynolds-averaged Navier-Stokes solutions on a parametric series of geometries which varied in leading-edge radius and/or circular-arc camber. The computational results showed that increasing leading-edge radius or camber can delay the onset of leading-edge separation on the leeside of a delta wing at a specific angle of attack. Reynolds number was varied from 1 x 10 to the 6th to 5 x 10 to the 6th for a turbulent boundary-layer and was shown to have a minor effect on the effectiveness of leading-edge radius and/or camber in delaying the onset of leading-edge separation. Both laminar and turbulent boundary-layer models were investigated at a Reynolds number of 1 x 10 to the 6th, and the predicted flow pattern was found to change from attached flow for the turbulent boundary-layer model to separated flow for the laminar boundary-layer model. Based upon these results, three wind-tunnel models have been designed to be tested in the Langley Unitary Plan Wind Tunnel.

Large-scale wind-tunnel investigation of a close-coupled canard-delta-wing fighter model through high angles of attack

NASA Technical Reports Server (NTRS)

Stoll, F.; Koenig, D. G.

1983-01-01

Data obtained through very high angles of attack from a large-scale, subsonic wind-tunnel test of a close-coupled canard-delta-wing fighter model are analyzed. The canard delays wing leading-edge vortex breakdown, even for angles of attack at which the canard is completely stalled. A vortex-lattice method was applied which gave good predictions of lift and pitching moment up to an angle of attack of about 20 deg, where vortex-breakdown effects on performance become significant. Pitch-control inputs generally retain full effectiveness up to the angle of attack of maximum lift, beyond which, effectiveness drops off rapidly. A high-angle-of-attack prediction method gives good estimates of lift and drag for the completely stalled aircraft. Roll asymmetry observed at zero sideslip is apparently caused by an asymmetry in the model support structure.

The effect of asymmetric vortex wake characteristics on a slender delta wing undergoing wing rock motion

NASA Technical Reports Server (NTRS)

Arena, A. S., Jr.; Nelson, R. C.

1989-01-01

An experimental investigation into the fluid mechanisms responsible for wing rock on a slender delta wing with 80 deg leading edge sweep has been conducted. Time history and flow visualization data are presented for a wide angle-of-attack range. The use of an air bearing spindle has allowed the motion of the wing to be free from bearing friction or mechanical hysteresis. A bistable static condition has been found in vortex breakdown at an angle of attack of 40 deg which causes an overshoot of the steady state rocking amplitude. Flow visualization experiments also reveal a difference in static and dynamic breakdown locations on the wing. A hysteresis loop in dynamic breakdown location similar to that seen on pitching delta wings was observed as the wing was undergoing the limit cycle oscillation.

Development of a High-Performance Fin-and-Tube Heat Exchanger with Vortex Generators for a Vending Machine

NASA Astrophysics Data System (ADS)

Iwasaki, Masamichi; Saito, Hiroshi; Mochizuki, Sadanari; Murata, Akira

The effect of delta-wing-vortex generators (combination of a delta wing and a delta winglet pair) on the heat transfer performance of fin-and-tube heat exchangers for vending machines has been investegated. Flow visualizations, numerical simulations and heat transfer experiments were conducted to find an optimum geometrical shape and arrangement of the vortex generators. Maximum heat transfer enhancement was achieved by the combination of (a) the delta wing with the apex angle of 86 degrees and (b) the delta winglet pair with the inline angle of 45 degrees. In relatively low Reynolds number range, about 40 % increase in heat transfer coefficient was attained with the above mentioned combination of the vortex generators compared to the ordinary heat exchangers with plain fins. It was revealed that the heat transfer enhancement was attributed to (1) the longitudinal vortexes generated by the delta wing and (2) the reduction of wake area behind the tube. It was also found that an increase in the apex angle of the delta wing brought about heat transfer enhancement, and the scale as well as the streggth of the induced longitudinal vortices played an important role in the heat transfer performance.

The DELTA MONSTER: An RPV designed to investigate the aerodynamics of a delta wing platform

NASA Technical Reports Server (NTRS)

Connolly, Kristen; Flynn, Mike; Gallagher, Randy; Greek, Chris; Kozlowski, Marc; Mcdonald, Brian; Mckenna, Matt; Sellar, Rich; Shearon, Andy

1989-01-01

The mission requirements for the performance of aerodynamic tests on a delta wind planform posed some problems, these include aerodynamic interference; structural support; data acquisition and transmission instrumentation; aircraft stability and control; and propulsion implementation. To eliminate the problems of wall interference, free stream turbulence, and the difficulty of achieving dynamic similarity between the test and actual flight aircraft that are associated with aerodynamic testing in wind tunnels, the concept of the remotely piloted vehicle which can perform a basic aerodynamic study on a delta wing was the main objective for the Green Mission - the Delta Monster. The basic aerodynamic studies were performed on a delta wing with a sweep angle greater than 45 degrees. These tests were performed at various angles of attack and Reynolds numbers. The delta wing was instrumented to determine the primary leading edge vortex formation and location, using pressure measurements and/or flow visualization. A data acquisition system was provided to collect all necessary data.

Investigation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Covell, Peter F.; Wood, Richard M.; Miller, David S.

1987-01-01

An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.

Navier-Stokes, dynamics and aeroelastic computations for vortical flows, buffet and flutter applications

NASA Technical Reports Server (NTRS)

Kandil, Osama A.

1993-01-01

Research on Navier-Stokes, dynamics, and aeroelastic computations for vortical flows, buffet, and flutter applications was performed. Progress during the period from 1 Oct. 1992 to 30 Sep. 1993 is included. Papers on the following topics are included: vertical tail buffet in vortex breakdown flows; simulation of tail buffet using delta wing-vertical tail configuration; shock-vortex interaction over a 65-degree delta wing in transonic flow; supersonic vortex breakdown over a delta wing in transonic flow; and prediction and control of slender wing rock.

An exploratory study of apex fence flaps on a 74 deg delta wing

NASA Technical Reports Server (NTRS)

Wahls, R. A.; Vess, R. J.

1985-01-01

An exploratory wind tunnel investigation was performed to observe the flow field effects produced by vertically deployed apex fences on a planar 74 degree delta wing. The delta shaped fences, each comprising approximately 3.375 percent of the wing area, were affixed along the first 25 percent of the wing leading edge in symmetric as well as asymmetric (i.e., fence on one side only) arrangements. The vortex flow field was visualized at angles of attack from 0 to 20 degrees using helium bubble and oil flow techniques; upper surface pressures were also measured along spanwise rows. The results were used to construct a preliminary description of the vortex patterns and induced pressures associated with vertical apex fence deployment. The objective was to obtain an initial evaluation of the potential of apex fences as vortex devices for subsonic lift modulation as well as lateral directional control of delta wing aircraft.

Similitude relations for buffet and wing rock on delta wings

NASA Astrophysics Data System (ADS)

Mabey, D. G.

1997-08-01

Vortex flow phenomena at high angles of incidence are of great interest to the designers of advanced combat aircraft. The steady phenomena (such as steady lift and pitching moments) are understood fairly well, whereas the unsteady phenomena are still uncertain. This paper addresses two important unsteady phenomena on delta wings. With regard to the frequency parameter of the quasi-periodic excitation caused by vortex bursting, a new correlation is established covering a range of sweep back from 60 to 75°. With regard to the much lower frequency parameter of limit-cycle rigid-body wing-rock, a new experiment shows conclusively that although the motion is non-linear, the frequency parameter can be predicted by quasi-steady theory. As a consequence, for a given sweep angle, the frequency parameter is inversely proportional to the square root of the inertia in roll. This is an important observation when attempting to extrapolate from model tests in wind tunnels to predict the wing-rock characteristics of aircraft.

Analytical observations on the aerodynamics of a delta wing with leading edge flaps

NASA Technical Reports Server (NTRS)

Oh, S.; Tavella, D.

1986-01-01

The effect of a leading edge flap on the aerodynamics of a low aspect ratio delta wing is studied analytically. The separated flow field about the wing is represented by a simple vortex model composed of a conical straight vortex sheet and a concentrated vortex. The analysis is carried out in the cross flow plane by mapping the wing trace, by means of the Schwarz-Christoffel transformation into the real axis of the transformed plane. Particular attention is given to the influence of the angle of attack and flap deflection angle on lift and drag forces. Both lift and drag decrease with flap deflection, while the lift-to-drag ratioe increases. A simple coordinate transformation is used to obtain a closed form expression for the lift-to-drag ratio as a function of flap deflection. The main effect of leading edge flap deflection is a partial suppression of the separated flow on the leeside of the wing. Qualitative comparison with experiments is presented, showing agreement in the general trends.

Effect of Thickness-to-Chord Ratio on Flow Structure of Low Swept Delta Wing

NASA Astrophysics Data System (ADS)

Gulsacan, Burak; Sencan, Gizem; Yavuz, Mehmet Metin

2017-11-01

The effect of thickness-to-chord (t/C) ratio on flow structure of a delta wing with sweep angle of 35 degree is characterized in a low speed wind tunnel using laser illuminated smoke visualization, particle image velocimetry, and surface pressure measurements. Four different t/C ratio varying from 4.75% to 19% are tested at angles of attack 4, 6, 8, and 10 degrees for Reynolds numbers Re =10,000 and 35,000. The results indicate that the effect of thickness-to-chord ratio on flow structure is quite substantial, such that, as the wing thickness increases, the flow structure transforms from leading edge vortex to three-dimensional separated flow regime. The wing with low t/C ratio of 4.75% experiences pronounced surface separation at significantly higher angle of attack compared to the wing with high t/C ratio. The results might explain some of the discrepancies reported in previously conducted studies related to delta wings. In addition, it is observed that the thickness of the shear layer separated from windward side of the wing is directly correlated with the thickness of the wing. To conclude, the flow structure on low swept delta wing is highly affected by t/C ratio, which in turn might indicate the potential usage of wing thickness as an effective flow control parameter.

Physical mechanisms of longitudinal vortexes formation, appearance of zones with high heat fluxes and early transition in hypersonic flow over delta wing with blunted leading edges

NASA Astrophysics Data System (ADS)

Alexandrov, S. V.; Vaganov, A. V.; Shalaev, V. I.

2016-10-01

Processes of vortex structures formation and they interactions with the boundary layer in the hypersonic flow over delta wing with blunted leading edges are analyzed on the base of experimental investigations and numerical solutions of Navier-Stokes equations. Physical mechanisms of longitudinal vortexes formation, appearance of abnormal zones with high heat fluxes and early laminar turbulent transition are studied. These phenomena were observed in many high-speed wind tunnel experiments; however they were understood only using the detailed analysis of numerical modeling results with the high resolution. Presented results allowed explaining experimental phenomena. ANSYS CFX code (the DAFE MIPT license) on the grid with 50 million nodes was used for the numerical modeling. The numerical method was verified by comparison calculated heat flux distributions on the wing surface with experimental data.

Reactive Flow Control of Delta Wing Vortex (Postprint)

DTIC Science & Technology

2006-08-01

wing aircraft. A substantial amount of research has been dedicated to the control of aerodynamic flows using both passive and active control mechanisms...Passive vortex control devices such as vortex generators and winglets attach to the wing and require no energy input. Passive vortex control...leading edges is also effective for changing the aerodynamic characteristics of delta wings [2] [3]. Gutmark and Guillot [5] proposed controlling

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 2; Small-Radius Leading Edge

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg. delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 84 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 4: Large-radius leading edge

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6) and 60 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Vol. 3: Medium-radius leading edge

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 120 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at Reynolds numbers of 6 x 10(exp 6), 60 x 10(exp 6), and 120 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Experimental Surface Pressure Data Obtained on 65 deg Delta Wing Across Reynolds Number and Mach Number Ranges. Volume 1; Sharp Leading Edge; [conducted in the Langley National Transonic Facility (NTF)

NASA Technical Reports Server (NTRS)

Chu, Julio; Luckring, James M.

1996-01-01

An experimental wind tunnel test of a 65 deg delta wing model with interchangeable leading edges was conducted in the Langley National Transonic Facility (NTF). The objective was to investigate the effects of Reynolds and Mach numbers on slender-wing leading-edge vortex flows with four values of wing leading-edge bluntness. Experimentally obtained pressure data are presented without analysis in tabulated and graphical formats across a Reynolds number range of 6 x 10(exp 6) to 36 x 10(exp 6) at a Mach number of 0.85 and across a Mach number range of 0.4 to 0.9 at a Reynolds number of 6 x 10(exp 6). Normal-force and pitching-moment coefficient plots for these Reynolds number and Mach number ranges are also presented.

Wind-tunnel investigation of a large-scale VTOL aircraft model with wing root and wing thrust augmentors. [Ames 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Aoyagi, K.; Aiken, T. N.

1979-01-01

Tests were conducted in the Ames 40 by 80 foot wind tunnel to determine the aerodynamic characteristics of a large-scale V/STOL aircraft model with thrust augmentors. The model had a double-delta wing of aspect ratio 1.65 with augmentors located in the wing root and the wing trailing edge. The supply air for the augmentor primary nozzles was provided by the YJ-97 turbojet engine. The airflow was apportioned approximately 74 percent to the wing root augmentor and 24 percent to wing augmentor. Results were obtained at several trailing-edge flap deflections with the nozzle jet-momentum coefficients ranging from 0 to 7.9. Three-component longitudinal data are presented with the agumentor operating with and without the horizontal tail. A limited amount of six component data are also presented.

Unsteady aerodynamic analysis of space shuttle vehicles. Part 2: Steady and unsteady aerodynamics of sharp-edged delta wings

NASA Technical Reports Server (NTRS)

Ericsson, L. E.; Reding, J. P.

1973-01-01

An analysis of the steady and unsteady aerodynamics of sharp-edged slender wings has been performed. The results show that slender wing theory can be modified to give the potential flow static and dynamic characteristics in incompressible flow. A semiempirical approximation is developed for the vortex-induced loads, and it is shown that the analytic approximation for sharp-edged slender wings gives good prediction of experimentally determined steady and unsteady aerodynamics at M = 0 and M = 1. The predictions are good not only for delta wings but also for so-called arrow and diamond wings. The results indicate that the effects of delta planform lifting surfaces can be included in a simple manner when determining elastic launch vehicle dynamic characteristics. For Part 1 see (N73-32763).

Navier-Stokes and Euler solutions for lee-side flows over supersonic delta wings. A correlation with experiment

NASA Technical Reports Server (NTRS)

Mcmillin, S. Naomi; Thomas, James L.; Murman, Earll M.

1990-01-01

An Euler flow solver and a thin layer Navier-Stokes flow solver were used to numerically simulate the supersonic leeside flow fields over delta wings which were observed experimentally. Three delta wings with 75, 67.5, and 60 deg leading edge sweeps were computed over an angle-of-attack range of 4 to 20 deg at a Mach number 2.8. The Euler code and Navier-Stokes code predict equally well the primary flow structure where the flow is expected to be separated or attached at the leading edge based on the Stanbrook-Squire boundary. The Navier-Stokes code is capable of predicting both the primary and the secondary flow features for the parameter range investigated. For those flow conditions where the Euler code did not predict the correct type of primary flow structure, the Navier-Stokes code illustrated that the flow structure is sensitive to boundary layer model. In general, the laminar Navier-Stokes solutions agreed better with the experimental data, especially for the lower sweep delta wings. The computational results and a detailed re-examination of the experimental data resulted in a refinement of the flow classifications. This refinement in the flow classification results in the separation bubble with the shock flow type as the intermediate flow pattern between separated and attached flows.

Wing kinematics and flexibility for optimal manoeuvring and escape

NASA Astrophysics Data System (ADS)

Wong, Jaime Gustav

Understanding how animals control the dynamic stall vortices in their wake is critical to developing micro-aerial vehicles and autonomous underwater vehicles, not to mention wind turbines, delta wings, and rotor craft that undergo similar dynamic stall processes. Applying this knowledge to biomimetic engineering problems requires progress in three areas: (i) understanding the flow physics of natural swimmers and flyers; (ii) developing flow measurement techniques to resolve this physics; and (iii) deriving low-cost models suitable for studying the vast parameter space observed in nature. This body of work, which consists of five research chapters, focuses on the leading-edge vortex (LEV) that forms on profiles undergoing rapid manoeuvres, delta wings, and similar devices. Lagrangian particle tracking is used throughout this thesis to track the mass and circulation transport in the LEV on manoeuvring profiles. The growth and development of the LEV is studied in relation to: flapping and plunging profile kinematics; spanwise flow from profile sweep and spanwise profile bending; and varying the angle-of-attack gradient along the profile span. Finally, scaling relationships derived from the observations above are used to develop a low-cost model for LEV growth, that is validated on a flat-plate delta wing. Together these results contribute to each of the three topics identified above, as a step towards developing robust, agile biomimetic swimmers and flyers.

Vortical structures of supersonic flow over a delta-wing on a flat plate

NASA Astrophysics Data System (ADS)

Wang, D. P.; Xia, Z. X.; Zhao, Y. X.; Wang, Q. H.; Liu, B.

2013-02-01

Employing the nanoparticle-based planar laser scattering (NPLS), supersonic flow over a delta-winged vortex generator on a flat plate was experimentally investigated in a supersonic quiet wind tunnel at Ma = 2.68. The fine structures of the flow field, shock waves, separation vortices, wake, and boundary layer transition were observed in the NPLS images. According to the time-correlation of the NPLS images and the measurement results of particle image velocimetry, the structural model of the flow field was improved further, and coherent wake structures were observed, which is of significance theoretically and in engineering application.

Conical Euler analysis and active roll suppression for unsteady vortical flows about rolling delta wings

NASA Technical Reports Server (NTRS)

Lee-Rausch, Elizabeth M.; Batina, John T.

1993-01-01

A conical Euler code was developed to study unsteady vortex-dominated flows about rolling, highly swept delta wings undergoing either forced motions or free-to-roll motions that include active roll suppression. The flow solver of the code involves a multistage, Runge-Kutta time-stepping scheme that uses a cell-centered, finite-volume, spatial discretization of the Euler equations on an unstructured grid of triangles. The code allows for the additional analysis of the free to-roll case by simultaneously integrating in time the rigid-body equation of motion with the governing flow equations. Results are presented for a delta wing with a 75 deg swept, sharp leading edge at a free-stream Mach number of 1.2 and at 10 deg, 20 deg, and 30 deg angle of attack alpha. At the lower angles of attack (10 and 20 deg), forced-harmonic analyses indicate that the rolling-moment coefficients provide a positive damping, which is verified by free-to-roll calculations. In contrast, at the higher angle of attack (30 deg), a forced-harmonic analysis indicates that the rolling-moment coefficient provides negative damping at the small roll amplitudes. A free-to-roll calculation for this case produces an initially divergent response, but as the amplitude of motion grows with time, the response transitions to a wing-rock type of limit cycle oscillation, which is characteristic of highly swept delta wings. This limit cycle oscillation may be actively suppressed through the use of a rate-feedback control law and antisymmetrically deflected leading-edge flaps. Descriptions of the conical Euler flow solver and the free-to roll analysis are included in this report. Results are presented that demonstrate how the systematic analysis of the forced response of the delta wing can be used to predict the stable, neutrally stable, and unstable free response of the delta wing. These results also give insight into the flow physics associated with unsteady vortical flows about delta wings undergoing forced motions and free-to-roll motions, including the active suppression of the wing-rock type phenomenon. The conical Euler methodology developed is directly extend able to three-dimensional calculations.

Conical Euler solution for a highly-swept delta wing undergoing wing-rock motion

NASA Technical Reports Server (NTRS)

Lee, Elizabeth M.; Batina, John T.

1990-01-01

Modifications to an unsteady conical Euler code for the free-to-roll analysis of highly-swept delta wings are described. The modifications involve the addition of the rolling rigid-body equation of motion for its simultaneous time-integration with the governing flow equations. The flow solver utilized in the Euler code includes a multistage Runge-Kutta time-stepping scheme which uses a finite-volume spatial discretization on an unstructured mesh made up of triangles. Steady and unsteady results are presented for a 75 deg swept delta wing at a freestream Mach number of 1.2 and an angle of attack of 30 deg. The unsteady results consist of forced harmonic and free-to-roll calculations. The free-to-roll case exhibits a wing rock response produced by unsteady aerodynamics consistent with the aerodynamics of the forced harmonic results. Similarities are shown with a wing-rock time history from a low-speed wind tunnel test.

An overview of selected NASP aeroelastic studies at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Spain, Charles V.; Soistmann, David L.; Parker, Ellen C.; Gibbons, Michael D.; Gilbert, Michael G.

1990-01-01

Following an initial discussion of the NASP flight environment, the results of recent aeroelastic testing of NASP-type highly swept delta-wing models in Langley's Transonic Dynamics Tunnel (TDT) are summarized. Subsonic and transonic flutter characteristics of a variety of these models are described, and several analytical codes used to predict flutter of these models are evaluated. These codes generally provide good, but conservative predictions of subsonic and transonic flutter. Also, test results are presented on a nonlinear transonic phenomena known as aileron buzz which occurred in the wind tunnel on highly swept delta wings with full-span ailerons. An analytical procedure which assesses the effects of hypersonic heating on aeroelastic instabilities (aerothermoelasticity) is also described. This procedure accurately predicted flutter of a heated aluminum wing on which experimental data exists. Results are presented on the application of this method to calculate the flutter characteristics of a fine-element model of a generic NASP configuration. Finally, it is demonstrated analytically that active controls can be employed to improve the aeroelastic stability and ride quality of a generic NASP vehicle flying at hypersonic speeds.

Aerodynamic performance of a wing with a deflected tip-mounted reverse half-delta wing

NASA Astrophysics Data System (ADS)

Lee, T.; Su, Y. Y.

2012-11-01

The impact of a tip-mounted 65°-sweep reverse half-delta wing (RHDW), set at different deflections, on the aerodynamic performance of a rectangular NACA 0012 wing was investigated experimentally at Re = 2.45 × 105. This study is a continuation of the work of Lee and Su (Exp Fluids 52(6):1593-1609, 2012) on the passive control of wing tip vortex by the use of a reverse half-delta wing. The present results show that for RHDW deflection with -5° ≤ δ ≤ +15°, the lift was found to increase nonlinearly with increasing δ compared to the baseline wing. The lift increment was accompanied by an increased total drag. For negative RHDW deflection with δ < -5°, the RHDW-induced lift decrement was, however, accompanied by an improved drag. The deflected RHDW also significantly modified and weakened the tip vortex, leading to a persistently lowered lift-induced drag, regardless of its deflection, compared to the baseline wing. Physical mechanisms responsible for the observed RHDW-induced phenomenon were also discussed.

Investigation of dynamic ground effect

NASA Technical Reports Server (NTRS)

Chang, Ray Chung; Muirhead, Vincent U.

1987-01-01

An experimental investigation of dynamic ground effect was conducted in the Univ. of Kansas wind tunnel using delta wings of 60, 70, 75 deg sweep; the XB-70 wing; and the F-104A wing. Both static and dynamic tests were made. Test data were compared to other test data, including dynamic flight test data of the XB-70 and F-104A. Limited flow visualization test were conducted. A significant dynamic effect was found for highly swept delta wings.

Evaluation of Blended Wing-Body Combinations with Curved Plan Forms at Mach Numbers Up to 3.50

NASA Technical Reports Server (NTRS)

Holdaway, George H.; Mellenthin, Jack A.

1960-01-01

This investigation is a continuation of the experimental and theoretical evaluation of the effects of wing plan-form variations on the aerodynamic performance characteristics of blended wing-body combinations. The present report compares previously tested straight-edged delta and arrow models which have leading-edge sweeps of 59.04 and 70-82 deg., respectively, with related models which have plan forms with curved leading and trailing edges designed to result in the same average sweeps in each case. All the models were symmetrical, without camber, and were generally similar having the same span, length, and aspect ratios. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local wing chords in a streamwise direction. The wing sections were computed by varying their shapes along with the body radii (blending process) to match the selected area distribution and the given plan form. The models were tested with transition fixed at Reynolds numbers of roughly 4,000,000 to 9,000,000, based on the mean aerodynamic chord of the wing. The characteristic effect of the wing curvature of the delta and arrow models was an increase at subsonic and transonic speeds in the lift-curve slopes which was partially reflected in increased maximum lift-drag ratios. Curved edges were not evaluated on a diamond plan form because a preliminary investigation indicated that the curvature considered would increase the supersonic zero-lift wave drag. However, after the test program was completed, a suitable modification for the diamond plan form was discovered. The analysis presented in the appendix indicates that large reductions in the zero-lift wave drag would be obtained at supersonic Mach numbers if the leading- and trailing-edge sweeps are made to differ by indenting the trailing edge and extending the root of the leading edge.

Recent Loads Calibration Experience With a Delta Wing Airplane

NASA Technical Reports Server (NTRS)

Jenkins, Jerald M.; Kuhl, Albert E.

1977-01-01

Aircraft which are designed for supersonic and hypersonic flight are evolving with delta wing configurations. An integral part of the evolution of all new aircraft is the flight test phase. Included in the flight test phase is an effort to identify and evaluate the loads environment of the aircraft. The most effective way of examining the loads environment is to utilize calibrated strain gages to provide load magnitudes. Using strain gage data to accomplish this has turned out to be anything but a straightforward task. The delta wing configuration has turned out to be a very difficult type of wing structure to calibrate. Elevated structural temperatures result in thermal effects which contaminate strain gage data being used to deduce flight loads. The concept of thermally calibrating a strain gage system is an approach to solving this problem. This paper will address how these problems were approached on a program directed toward measuring loads on the wing of a large, flexible supersonic aircraft. Structural configurations typical of high-speed delta wing aircraft will be examined. The temperature environment will be examined to see how it induces thermal stresses which subsequently cause errors in loads equations used to deduce the flight loads.

A Wind-Tunnel Investigation of the Development of Lift on Wings in Accelerated Longitudinal Motion

NASA Technical Reports Server (NTRS)

Turner, Thomas R.

1960-01-01

An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the development of lift on a wing during a simulated constant-acceleration catapult take-off. The investigation included models of a two-dimensional wing, an unswept wing having an aspect ratio of 6, a 35 deg. swept wing having an aspect ratio of 3.05, and a 60 deg. delta wing having an aspect ratio of 2.31. All the wings investigated developed at least 90 percent of their steady-state lift in the first 7 chord lengths of travel. The development of lift was essentially independent of the acceleration when based on chord lengths traveled, and was in qualitative agreement with theory.

Steady and unsteady transonic pressure measurements on a clipped delta wing for pitching and control-surface oscillations

NASA Technical Reports Server (NTRS)

Hess, Robert W.; Cazier, F. W., Jr.; Wynne, Eleanor C.

1986-01-01

Steady and unsteady pressures were measured on a clipped delta wing with a 6-percent circular-arc airfoil section and a leading-edge sweep angle of 50.40 deg. The model was oscillated in pitch and had an oscillating trailing-edge control surface. Measurements were concentrated over a Mach number range from 0.88 to 0.94; less extensive measurements were made at Mach numbers of 0.40, 0.96, and 1.12. The Reynolds number based on mean chord was approximately 10 x 10 to the 6th power. The interaction of wing or control-surface deflection with the formation of shock waves and with a leading-edge vortex generated complex pressure distributions that were sensitive to frequency and to small changes in Mach number at transonic speeds.

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Flight of Sharovipteryx mirabilis: the world's first delta-winged glider.

PubMed

Dyke, G J; Nudds, R L; Rayner, J M V

2006-07-01

The 225 million-year-old reptile Sharovipteryx mirabilis was the world's first delta-winged glider; this remarkable animal had a flight surface composed entirely of a hind-limb membrane. We use standard delta-wing aerodynamics to reconstruct the flight of S. mirabilis demonstrating that wing shape could have been controlled simply by protraction of the femora at the knees, and by variation in incidence of a small forelimb canard. Our method has allowed us to address the question of how identifying realistic glide performance can be used to set limits on aerodynamic design in this small animal. Our novel interpretation of the bizarre flight mode of S. mirabilis is the first based directly on interpretation of the fossil itself and the first grounded in aerodynamics.

Euler and Potential Experiment/CFD Correlations for a Transport and Two Delta-Wing Configurations

NASA Technical Reports Server (NTRS)

Hicks, R. M.; Cliff, S. E.; Melton, J. E.; Langhi, R. G.; Goodsell, A. M.; Robertson, D. D.; Moyer, S. A.

1990-01-01

A selection of successes and failures of Computational Fluid Dynamics (CFD) is discussed. Experiment/CFD correlations involving full potential and Euler computations of the aerodynamic characteristics of four commercial transport wings and two low aspect ratio, delta wing configurations are shown. The examples consist of experiment/CFD comparisons for aerodynamic forces, moments, and pressures. Navier-Stokes equations are not considered.

Flow-field in a vortex with breakdown above sharp edged delta wings

NASA Technical Reports Server (NTRS)

Hayashi, Y.; Nakaya, T.

1978-01-01

The behavior of vortex-flow, accompanied with breakdown, formed above sharp-edged delta wings, was studied experimentally as well as theoretically. Emphasis is placed particularly on the criterion for the breakdown at sufficiently large Reynolds numbers

Aerodynamic Interaction between Delta Wing and Hemisphere-Cylinder in Supersonic Flow

NASA Astrophysics Data System (ADS)

Nishino, Atsuhiro; Ishikawa, Takahumi; Nakamura, Yoshiaki

As future space vehicles, Reusable Launch Vehicle (RLV) needs to be developed, where there are two kinds of RLV: Single Stage To Orbit (SSTO) and Two Stage To Orbit (TSTO). In the latter case, the shock/shock interaction and shock/boundary layer interaction play a key role. In the present study, we focus on the supersonic flow field with aerodynamic interaction between a delta wing and a hemisphere-cylinder, which imitate a TSTO, where the clearance, h, between the delta wing and hemisphere-cylinder is a key parameter. As a result, complicated flow patterns were made clear, including separation bubbles.

System Identification of a Vortex Lattice Aerodynamic Model

NASA Technical Reports Server (NTRS)

Juang, Jer-Nan; Kholodar, Denis; Dowell, Earl H.

2001-01-01

The state-space presentation of an aerodynamic vortex model is considered from a classical and system identification perspective. Using an aerodynamic vortex model as a numerical simulator of a wing tunnel experiment, both full state and limited state data or measurements are considered. Two possible approaches for system identification are presented and modal controllability and observability are also considered. The theory then is applied to the system identification of a flow over an aerodynamic delta wing and typical results are presented.

Wind tunnel investigation of nacelle-airframe interference at Mach numbers of 0.9 to 1.4 - pressure data, volume 1

NASA Technical Reports Server (NTRS)

Bencze, D. P.

1976-01-01

Detailed interference force and pressure data were obtained on a representative wing-body nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The model was mounted on a six component force balance, and the left hand wing was pressure instrumented. Each of the two right hand nacelles was mounted on a six component force balance housed in the thickness of the nacelle, while each of the left hand nacelles was pressure instrumented. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass flow ratio. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components.

Investigation at High Subsonic Speeds of the Static Longitudinal and Lateral Stability Characteristics of Two Canard Airplane Configurations

NASA Technical Reports Server (NTRS)

Sleeman, William C., Jr.

1957-01-01

The present investigation was conducted in the Langley high-speed 7-by 10-foot tunnel to determine the static longitudinal and lateral stability characteristics at high subsonic speeds of two canard airplane configurations previously tested at supersonic speeds. The Mach number range of this investigation extended from 0.60 to 0.94 and a maximum angle-of-attack range of -2dewg to 24deg was obtained at the lowest test Mach number. Two wing plan forms of equal area were studied in the present tests; one was a 60deg delta wing and the other was a trapezoid wing having an aspect ratio of 3, taper ratio of 0.143, and an unswept 80-percent-chord line. The canard control had a trapezoidal plan form and its area was approximately 11.5 percent of the wing area. The model also had a low-aspect-ratio highly swept vertical tail and twin ventral fins. The longitudinal control characteristics of the models were consistent with past experience at low speed on canard configurations in that stalling of the canard surface occurred at moderate and high control deflections for moderate values of angle of attack. This stalling could impose appreciable limitations on the maximum trim-lift coefficient attainable. The control effectiveness and maximum value of trim-lift was significantly increased by addition of a body flap having a conical shape and located slightly behind the canard surface on the bottom of the body. Addition of the canard surface at 0deg deflection had relatively little effect on overall directional stability of the delta-wing configuration; however, deflection of the canard surface from 0deg to 10deg had a large favorable effect on directional stability at high angles of attack for both the trapezoid- and delta-wing configurations.

A study of roll attractor and wing rock of delta wings at high angles of attack

NASA Technical Reports Server (NTRS)

Niranjana, T.; Rao, D. M.; Pamadi, Bandu N.

1993-01-01

Wing rock is a high angle of attack dynamic phenomenon of limited cycle motion predominantly in roll. The wing rock is one of the limitations to combat effectiveness of the fighter aircraft. Roll Attractor is the steady state or equilibrium trim angle (phi(sub trim)) attained by the free-to-roll model, held at some angle of attack, and released form rest at a given initial roll (bank) angle (phi(sub O)). Multiple roll attractors are attained at different trim angles depending on initial roll angle. The test facility (Vigyan's low speed wind tunnel) and experimental work is presented here along with mathematical modelling of roll attractor phenomenon and analysis and comparison of predictions with experimental data.

Flow Behavior in Side-View Plane of Pitching Delta Wing

NASA Astrophysics Data System (ADS)

Pektas, Mehmet Can; Tasci, Mehmet Oguz; Karasu, Ilyas; Sahin, Besir; Akilli, Huseyin

2018-06-01

In the present investigation, a delta wing which has 70° sweep angle, Λ was oscillated on its midcord according to the equation of α(t)=αm+α0sin(ωet). This study focused on understanding the effect of pitching and characterizing the interaction of vortex breakdown with oscillating leading edges under different yaw angles, β over a slender delta wing. The value of mean angle of attack, αm was taken as 25°. The yaw angle, β was varied with an interval of 4° over the range of 0°≤β≤ 16°. The delta wing was sinusoidally pitched within the range of period of time 5s≤Te≤60s and reduced frequency was set as K=0.16, 0.25, 0.49, 1.96 and lastly amplitude of pitching motion was arranged as α0=±5°.Formations and locations of vortex breakdown were investigated by using the dye visualization technique in side view plane.

An experimental investigation of vortex breakdown on a delta wing

NASA Technical Reports Server (NTRS)

Payne, F. M.; Nelson, R. C.

1986-01-01

An experimental investigation of vortex breakdown on delta wings at high angles is presented. Thin delta wings having sweep angles of 70, 75, 80 and 85 degrees are being studied. Smoke flow visualization and the laser light sheet technique are being used to obtain cross-sectional views of the leading edge vortices as they break down. At low tunnel speeds (as low as 3 m/s) details of the flow, which are usually imperceptible or blurred at higher speeds, can be clearly seen. A combination of lateral and longitudinal cross-sectional views provides information on the three dimensional nature of the vortex structure before, during and after breakdown. Whereas details of the flow are identified in still photographs, the dynamic characteristics of the breakdown process were recorded using high speed movies. Velocity measurements were obtained using a laser Doppler anemometer with the 70 degree delta wing at 30 degrees angle of attack. The measurements show that when breakdown occurs the core flow transforms from a jet-like flow to a wake-like flow.

Leading edge vortex control on a delta wing with dielectric barrier discharge plasma actuators

NASA Astrophysics Data System (ADS)

Shen, Lu; Wen, Chih-yung

2017-06-01

This paper presents an experimental investigation of the application of dielectric barrier discharge (DBD) plasma actuators on a slender delta wing to control the leading edge vortices (LEVs). The experiments are conducted in a wind tunnel with a Reynolds number of 50 000 based on the chord length. The smoke flow visualization reveals that the DBD plasma actuators at the leading edges significantly modify the vortical flow structure over the delta wing. It is noted that symmetric control at both semi-spans and asymmetric control at a single semi-span leads to opposite effects on the local LEVs. Particle image velocimetry (PIV) indicates that the shear layer is deformed by the actuators. Therefore, both the strength and the shape of the LEV cores are deeply affected. The six-component force measurement shows that the DBD plasma actuators have a limited effect on lift and drag while inducing relatively large moments. This suggests that the DBD plasma actuator is a promising technique for delta wing maneuvering.

A vortex-filament and core model for wings with edge vortex separation

NASA Technical Reports Server (NTRS)

Pao, J. L.; Lan, C. E.

1981-01-01

A method for predicting aerodynamic characteristics of slender wings with edge vortex separation was developed. Semiempirical but simple methods were used to determine the initial positions of the free sheet and vortex core. Comparison with available data indicates that: the present method is generally accurate in predicting the lift and induced drag coefficients but the predicted pitching moment is too positive; the spanwise lifting pressure distributions estimated by the one vortex core solution of the present method are significantly better than the results of Mehrotra's method relative to the pressure peak values for the flat delta; the two vortex core system applied to the double delta and strake wing produce overall aerodynamic characteristics which have good agreement with data except for the pitching moment; and the computer time for the present method is about two thirds of that of Mehrotra's method.

A vortex-filament and core model for wings with edge vortex separation

NASA Technical Reports Server (NTRS)

Pao, J. L.; Lan, C. E.

1982-01-01

A vortex filament-vortex core method for predicting aerodynamic characteristics of slender wings with edge vortex separation was developed. Semi-empirical but simple methods were used to determine the initial positions of the free sheet and vortex core. Comparison with available data indicates that: (1) the present method is generally accurate in predicting the lift and induced drag coefficients but the predicted pitching moment is too positive; (2) the spanwise lifting pressure distributions estimated by the one vortex core solution of the present method are significantly better than the results of Mehrotra's method relative to the pressure peak values for the flat delta; (3) the two vortex core system applied to the double delta and strake wings produce overall aerodynamic characteristics which have good agreement with data except for the pitching moment; and (4) the computer time for the present method is about two thirds of that of Mehrotra's method.

Spacecraft boost and abort guidance and control systems requirement study, boost dynamics and control analysis study. Exhibit A: Boost dynamics and control anlaysis

NASA Technical Reports Server (NTRS)

Williams, F. E.; Price, J. B.; Lemon, R. S.

1972-01-01

The simulation developments for use in dynamics and control analysis during boost from liftoff to orbit insertion are reported. Also included are wind response studies of the NR-GD 161B/B9T delta wing booster/delta wing orbiter configuration, the MSC 036B/280 inch solid rocket motor configuration, the MSC 040A/L0X-propane liquid injection TVC configuration, the MSC 040C/dual solid rocket motor configuration, and the MSC 049/solid rocket motor configuration. All of the latest math models (rigid and flexible body) developed for the MSC/GD Space Shuttle Functional Simulator, are included.

Space shuttle: Effect of configuration changes on the directional characteristics of a GD/C booster Mach no. 1.2 - 4.96

NASA Technical Reports Server (NTRS)

Brickey, J.; Brice, T.; Marks, K. E.

1971-01-01

Force tests on a 0.0035-scale model of the General Dynamics/Convair aerospace space shuttle B-15B-1 booster were conducted in the MSFC trisonic wind tunnel. The configuration has a low delta wing, all-movable delta-planform canard controls, and a single vertical tail. The test was devoted to investigating the effects of various configuration variables upon lateral-directional characteristics. These variables included wing dihedral, rudder flare, and body flap deflection. Yaw runs were made at angles of attack of 6, 10, 15, 25, 30, and 35 degrees. The Mach number range for this test was 1.20 to 4.96.

Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

NASA Technical Reports Server (NTRS)

Hernandez, Gloria; Wood, Richard M.; Covell, Peter F.

1994-01-01

An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

Theoretical study of aerodynamic characteristics of wings having vortex flow

NASA Technical Reports Server (NTRS)

Reddy, C. S.

1979-01-01

The aerodynamic characteristics of slender wings having separation induced vortex flows are investigated by employing three different computer codes--free vortex sheet, quasi vortex lattice, and suction analogy methods. Their capabilities and limitations are examined, and modifications are discussed. Flat wings of different configurations: arrow, delta, and diamond shapes, as well as cambered delta wings, are studied. The effect of notch ratio on the load distributions and the longitudinal characteristics of a family of arrow and diamond wings is explored. The sectional lift coefficients and the accumulated span loadings are determined for an arrow wing and are seen to be unusual in comparison with the attached flow results. The theoretically predicted results are compared with the existing experimental values.

Exhaust Plume Effects on Sonic Boom for a Delta Wing and a Swept Wing-Body Model

NASA Technical Reports Server (NTRS)

Castner, Raymond; Lake, Troy

2012-01-01

Supersonic travel is not allowed over populated areas due to the disturbance caused by the sonic boom. Research has been performed on sonic boom reduction and has included the contribution of the exhaust nozzle plume. Plume effect on sonic boom has progressed from the study of isolated nozzles to a study with four exhaust plumes integrated with a wing-body vehicle. This report provides a baseline analysis of the generic wing-body vehicle to demonstrate the effect of the nozzle exhaust on the near-field pressure profile. Reductions occurred in the peak-to-peak magnitude of the pressure profile for a swept wing-body vehicle. The exhaust plumes also had a favorable effect as the nozzles were moved outward along the wing-span.

Pressure measurements on a thick cambered and twisted 58 deg delta wing at high subsonic speeds

NASA Technical Reports Server (NTRS)

Chu, Julio; Lamar, John E.

1987-01-01

A pressure experiment at high subsonic speeds was conducted by a cambered and twisted thick delta wing at the design condition (Mach number 0.80), as well as at nearby Mach numbers (0.75 and 0.83) and over an angle-of-attack range. Effects of twin vertical tails on the wing pressure measurements were also assessed. Comparisons of detailed theoretical and experimental surface pressures and sectional characteristics for the wing alone are presented. The theoretical codes employed are FLO-57, FLO-28, PAN AIR, and the Vortex Lattice Method-Suction Analogy.

Flow field over the wing of a delta-wing fighter model with vortex control devices at Mach 0.6 to 1.2

NASA Technical Reports Server (NTRS)

Bare, E. Ann; Reubush, David E.; Haddad, Raymond C.

1992-01-01

As part of a cooperative research program between NASA, McDonnell Douglas Corporation, and Wright Research and Development Center, a flow field investigation was conducted on a 7.52 percent scale windtunnel model of an advanced fighter aircraft design. The investigation was conducted in the Langley 16 ft Transonic Tunnel at Mach numbers of 0.6, 0.9, and 1.2. Angle of attack was varied from -4 degrees to 30 degrees and the model was tested at angles of sideslip of 0, 5, and -5 degrees. Data for the over the wing flow field were obtained at four axial survey stations by the use of six 5 hole conical probes mounted on a survey mechanism. The wing leading edge primary vortex exerted the greatest influence in terms of total pressure loss on the over the wing flow field in the area surveyed. A number of vortex control devices were also investigated. They included two different apex flaps, wing leading edge vortex flaps, and small large wing fences. The vortex flap and both apex flaps were beneficial in controlling the wing leading edge primary vortex.

Self-Excited Roll Oscillations of Non-Slender Wings

DTIC Science & Technology

2010-03-01

on low sweep delta wings ( Yavuz et al. 2004; Taylor and Gursul 2004). Seeding was provided by a smoke machine placed in the low- speed section of...NV. Yavuz , M.M., Elkhoury, M., Rockwell, D., 2004, “Near-surface topology and flow structure on a delta wing”, AIAA Journal, vol. 42, no. 2, pp

A wind-tunnel investigation at high subsonic speeds of the lateral control characteristics of various plain spoiler configurations on a 3-percent-thick 60 degree delta wing

NASA Technical Reports Server (NTRS)

Wiley, Harleth G

1954-01-01

Results are presented of wind-tunnel investigations at Mach numbers of 0.60 to 0.94 and angles of attack of -2 degrees to about 24 degrees to determine the lateral control characteristics of spoilers with various wing chord-wise and spanwise locations and spoiler spans and deflections on thin 60 degree delta wing of NACA 65a003 airfoil section parallel to free stream.

An experimental investigation of delta wing vortex flow with and without external jet blowing

NASA Technical Reports Server (NTRS)

Iwanski, Kenneth P.; Ng, T. Terry; Nelson, Robert C.

1989-01-01

A visual and quantitative study of the vortex flow field over a 70-deg delta wing with an external jet blowing parallel to and at the leading edge was conducted. In the experiment, the vortex core was visually marked with TiCl4, and LDA was used to measure the velocity parallel and normal to the wing surface. It is found that jet blowing moved vortex breakdown farther downstream from its natural position and influenced the breakdown characteristics.

Wind tunnel investigation of Nacelle-Airframe interference at Mach numbers of 0.9 to 1.4-pressure data, volume 2

NASA Technical Reports Server (NTRS)

Bencze, D. P.

1976-01-01

Detailed interference force and pressure data were obtained on a representative wing-body nacelle combination at Mach numbers of 0.9 to 1.4. The model consisted of a delta wing-body aerodynamic force model with four independently supported nacelles located beneath the wing-body combination. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass flow ratio. Four different configurations were tested to identify various interference forces and pressures on each component; these included tests of the isolated nacelle, the isolated wing-body combination, the four nacelles as a unit, and the total wing-body-nacelle combination. Nacelle axial location, relative to both the wing-body combination and to each other, was the most important variable in determining the net interference among the components. The overall interference effects were found to be essentially constant over the operating angle-of-attack range of the configuration, and nearly independent of nacelle mass flow ratio.

Control of Flow Structure on Non-Slender Delta Wing: Bio-inspired Edge Modifications, Passive Bleeding, and Pulsed Blowing

NASA Astrophysics Data System (ADS)

Yavuz, Mehmet Metin; Celik, Alper; Cetin, Cenk

2016-11-01

In the present study, different flow control approaches including bio-inspired edge modifications, passive bleeding, and pulsed blowing are introduced and applied for the flow over non-slender delta wing. Experiments are conducted in a low speed wind tunnel for a 45 degree swept delta wing using qualitative and quantitative measurement techniques including laser illuminated smoke visualization, particle image velocimety (PIV), and surface pressure measurements. For the bio-inspired edge modifications, the edges of the wing are modified to dolphin fluke geometry. In addition, the concept of flexion ratio, a ratio depending on the flexible length of animal propulsors such as wings, is introduced. For passive bleeding, directing the free stream air from the pressure side of the planform to the suction side of the wing is applied. For pulsed blowing, periodic air injection through the leading edge of the wing is performed in a square waveform with 25% duty cycle at different excitation frequencies and compared with the steady and no blowing cases. The results indicate that each control approach is quite effective in terms of altering the overall flow structure on the planform. However, the success level, considering the elimination of stall or delaying the vortex breakdown, depends on the parameters in each method.

Unsteady surface pressure measurements on a slender delta wing undergoing limit cycle wing rock

NASA Technical Reports Server (NTRS)

Arena, Andrew S., Jr.; Nelson, Robert C.

1991-01-01

An experimental investigation of slender wing limit cycle motion known as wing rock was investigated using two unique experimental systems. Dynamic roll moment measurements and visualization data on the leading edge vortices were obtained using a free to roll apparatus that incorporates an airbearing spindle. In addition, both static and unsteady surface pressure data was measured on the top and bottom surfaces of the model. To obtain the unsteady surface pressure data a new computer controller drive system was developed to accurately reproduce the free to roll time history motions. The data from these experiments include, roll angle time histories, vortex trajectory data on the position of the vortices relative to the model's surface, and surface pressure measurements as a function of roll angle when the model is stationary or undergoing a wing rock motion. The roll time history data was numerically differentiated to determine the dynamic roll moment coefficient. An analysis of these data revealed that the primary mechanism for the limit cycle behavior was a time lag in the position of the vortices normal to the wing surface.

78 FR 1133 - Noise Certification Standards for Tiltrotors

Federal Register 2010, 2011, 2012, 2013, 2014

2013-01-08

... the wing tips that vary in pitch from near vertical to near horizontal configuration relative to the wing and fuselage. Vertical takeoff and landing (VTOL) mode means the aircraft state or configuration... arithmetic sum of delta 1 and the term -7.5 log (QK/QrKr) from delta 2 must not in total exceed 2.0 EPNdB; (b...

Lee side flow for slender delta wings of finite thickness

NASA Technical Reports Server (NTRS)

Szodruch, J. G.

1980-01-01

An experimental and theoretical investigation carried out to determine the lee side flow field over delta wings at supersonic speeds is presented. A theoretical method to described the flow field is described, where boundary conditions as a result of the experimental study are needed. The computed flow field with shock induced separation is satisfactory.

Space shuttle: Aerodynamic stability and control characteristics of the NASA/MSC .006 scale 040-A delta wing orbiter

NASA Technical Reports Server (NTRS)

Click, P. L.; Michana, D. J.; Sarver, D. A.

1971-01-01

Experimental aerodynamic investigations were made on a .006 scale model 040-A delta wing space shuttle orbiter configuration. These tests were conducted to determine six-degree-of-freedom force and moment data for preliminary stability and control analysis. Data were obtained over a Mach number range from 0.6 to 4.96 at angles of attack from -10 deg to 50 deg at zero degrees sideslip and at angles of sideslip from -10 deg to 10 deg at constants angles of attack of 0 deg, 15 deg, 30 deg, and 45 deg. Various aileron, elevator, (elevon) rudder and rudder flare deflection angles were tested to establish the control effectiveness and vehicle stability. Model component buildup data were also obtained to provide a data base for future configuration modifications. Plotted data results are presented in both the body and stability axis system.

Wing planform effects at supersonic speeds for an advanced fighter configuration

NASA Technical Reports Server (NTRS)

Wood, R. M.; Miller, D. S.

1984-01-01

Four advanced fighter configurations, which differed in wing planform and airfoil shape, were investigated in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, and 2.16. Supersonic data were obtained on the four uncambered wings, which were each attached to a single fighter fuselage. The fuselage geometry varied in cross-sectional shape and had two side-mounted, flow-through, half-axisymmetric inlets. Twin vertical tails were attached to the fuselage. The four planforms tested were a 65 deg delta wing, a combination of a 20 deg trapezoidal wing and a 45 deg horizontal tail, a 70 deg/30 deg cranked wing, and a 70 deg/66 deg crank wing, where the angle values refer to the leading-edge sweep angle of the lifting-surface planform. Planform effects on a single fuselage representative of an advanced fighter aircraft were studied. Results show that the highly swept cranked wings exceeded the aerodynamic performance levels, at low lift coefficients, of the 65 deg delta wing and the 20 deg trapezoidal wing at trimmed and untrimmed conditions.

Experimental study of supersonic viscous leeside flow over a slender delta wing

NASA Technical Reports Server (NTRS)

Szodruch, J.

1980-01-01

An investigation was conducted to study in detail the vortical flow over the leeward side of a 70 deg swept delta wing having subsonic and supersonic leading edges. Two types of flow were encountered and studied, namely leading edge separation and separation with a shock. Especially for the latter type, Reynolds number plays an important role and unexpected strong streamwise vortices were observed. An optical method is described to obtain a first aproximation of shear stress values in the streamwise direction across the wing span.

Conical Euler simulation and active suppression of delta wing rocking motion

NASA Technical Reports Server (NTRS)

Lee, Elizabeth M.; Batina, John T.

1990-01-01

A conical Euler code was developed to study unsteady vortex-dominated flows about rolling highly-swept delta wings, undergoing either forced or free-to-roll motions including active roll suppression. The flow solver of the code involves a multistage Runge-Kutta time-stepping scheme which uses a finite volume spatial discretization of the Euler equations on an unstructured grid of triangles. The code allows for the additional analysis of the free-to-roll case, by including the rigid-body equation of motion for its simultaneous time integration with the governing flow equations. Results are presented for a 75 deg swept sharp leading edge delta wing at a freestream Mach number of 1.2 and at alpha equal to 10 and 30 deg angle of attack. A forced harmonic analysis indicates that the rolling moment coefficient provides: (1) a positive damping at the lower angle of attack equal to 10 deg, which is verified in a free-to-roll calculation; (2) a negative damping at the higher angle of attack equal to 30 deg at the small roll amplitudes. A free-to-roll calculation for the latter case produces an initially divergent response, but as the amplitude of motion grows with time, the response transitions to a wing-rock type of limit cycle oscillation. The wing rocking motion may be actively suppressed, however, through the use of a rate-feedback control law and antisymmetrically deflected leading edge flaps. The descriptions of the conical Euler flow solver and the free-to-roll analysis are presented. Results are also presented which give insight into the flow physics associated with unsteady vortical flows about forced and free-to-roll delta wings, including the active roll suppression of this wing-rock phenomenon.

75 FR 39814 - Airworthiness Directives; Airbus Model A330-200 and -300, and Model A340-200, -300, -500 and -600...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-07-13

... Delta Air Lines (DAL) supports the intent of the NPRM. Request To Refer to Updated Revisions of Service... work `on-wing,' operators using PALL Service Bulletin 6753-20-2, Issue 4, dated November 21, 2009, for...

75 FR 5056 - Taking and Importing Marine Mammals; Taking Marine Mammals Incidental to Space Vehicle and Test...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-02-01

... hereby given that a letter of authorization (LOA) has been issued to the 30th Space Wing, U.S. Air Force... vehicle programs use VAFB to launch satellites into polar orbit: Delta II, Taurus, Atlas V, Delta IV..., and fixed-wing aircrafts are launched from VAFB. The activities under these regulations create two...

76 FR 6448 - Taking and Importing Marine Mammals; Taking Marine Mammals Incidental to Space Vehicle and Test...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-02-04

..., notification is hereby given that a letter of authorization (LOA) has been issued to the 30th Space Wing, U.S... launch satellites into polar orbit: Delta II; Taurus; Atlas V; Delta IV; Falcon; and Minotaur. Also a variety of small missiles, several types of interceptor and target vehicles, and fixed-wing aircrafts are...

76 FR 36001 - Noise Certification Standards for Tiltrotors

Federal Register 2010, 2011, 2012, 2013, 2014

2011-06-21

... and landing, within the powered-lift category, with rotors mounted at or near the wing tips that vary in pitch from near vertical to near horizontal configuration relative to the wing and fuselage... arithmetic sum of delta 1 and the term -7.5 log (QK/QrKr) from delta 2 must not in total exceed 2.0 EPNdB; (b...

Control of Flow Structure on Low Swept Delta Wing with Steady Leading Edge Blowing

NASA Astrophysics Data System (ADS)

Ozturk, Ilhan; Zharfa, Mohammadreza; Yavuz, Mehmet Metin

2014-11-01

Interest in unmanned combat air vehicles (UCAVs) and micro air vehicles (MAVs) has stimulated investigation of the flow structure, as well as its control, on delta wings having low and moderate values of sweep angle. In the present study, the flow structure is characterized on a delta wing of low sweep 35-degree angle, which is subjected to steady leading edge blowing. The techniques of laser illuminated smoke visualization, laser Doppler anemometry (LDA), and surface pressure measurements are employed to investigate the steady and unsteady nature of the flow structure on delta wing, in relation to the dimensionless magnitude of the blowing coefficient. Using statistics and spectral analysis, unsteadiness of the flow structure is studied in detail. Different injection locations are utilized to apply different blowing patterns in order to identify the most efficient control, which provides the upmost change in the flow structure with the minimum energy input. The study aims to find the optimum flow control strategy to delay or to prevent the stall and possibly to reduce the buffeting on the wing surface. Since the blowing set-up is computer controlled, the unsteady blowing patterns compared to the present steady blowing patterns will be studied next. This project was supported by the Scientific and Technological Research Council of Turkey (Project Number: 3501 111M732).

Flight test results from a supercritical mission adaptive wing with smooth variable camber

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Webb, Lannie D.; Friend, Edward L.; Lokos, William A.

1992-01-01

The mission adaptive wing (MAW) consisted of leading- and trailing-edge variable-camber surfaces that could be deflected in flight to provide a near-ideal wing camber shape for any flight condition. These surfaces featured smooth, flexible upper surfaces and fully enclosed lower surfaces, distinguishing them from conventional flaps that have discontinuous surfaces and exposed or semiexposed mechanisms. Camber shape was controlled by either a manual or automatic flight control system. The wing and aircraft were extensively instrumented to evaluate the local flow characteristics and the total aircraft performance. This paper discusses the interrelationships between the wing pressure, buffet, boundary-layer and flight deflection measurement system analyses and describes the flight maneuvers used to obtain the data. The results are for a wing sweep of 26 deg, a Mach number of 0.85, leading and trailing-edge cambers (delta(sub LE/TE)) of 0/2 and 5/10, and angles of attack from 3.0 deg to 14.0 deg. For the well-behaved flow of the delta(sub LE/TE) = 0/2 camber, a typical cruise camber shape, the local and global data are in good agreement with respect to the flow properties of the wing. For the delta(sub LE/TE) = 5/10 camber, a maneuvering camber shape, the local and global data have similar trends and conclusions, but not the clear-cut agreement observed for cruise camber.

Flow field predictions for a slab delta wing at incidence

NASA Technical Reports Server (NTRS)

Conti, R. J.; Thomas, P. D.; Chou, Y. S.

1972-01-01

Theoretical results are presented for the structure of the hypersonic flow field of a blunt slab delta wing at moderately high angle of attack. Special attention is devoted to the interaction between the boundary layer and the inviscid entropy layer. The results are compared with experimental data. The three-dimensional inviscid flow is computed numerically by a marching finite difference method. Attention is concentrated on the windward side of the delta wing, where detailed comparisons are made with the data for shock shape and surface pressure distributions. Surface streamlines are generated, and used in the boundary layer analysis. The three-dimensional laminar boundary layer is computed numerically using a specially-developed technique based on small cross-flow in streamline coordinates. In the rear sections of the wing the boundary layer decreases drastically in the spanwise direction, so that it is still submerged in the entropy layer at the centerline, but surpasses it near the leading edge. Predicted heat transfer distributions are compared with experimental data.

Numerical simulation of a powered-lift landing, tracking flow features using overset grids, and simulation of high lift devices on a fighter-lift-and-control wing

NASA Technical Reports Server (NTRS)

Chawla, Kalpana

1993-01-01

Attached as appendices to this report are documents describing work performed on the simulation of a landing powered-lift delta wing, the tracking of flow features using overset grids, and the simulation of flaps on the Wright Patterson Lab's fighter-lift-and-control (FLAC) wing. Numerical simulation of a powered-lift landing includes the computation of flow about a delta wing at four fixed heights as well as a simulated landing, in which the delta wing descends toward the ground. Comparison of computed and experimental lift coefficients indicates that the simulations capture the qualitative trends in lift-loss encountered by thrust-vectoring aircraft operating in ground effect. Power spectra of temporal variations of pressure indicate computed vortex shedding frequencies close to the jet exit are in the experimentally observed frequency range; the power spectra of pressure also provide insights into the mechanisms of lift oscillations. Also, a method for using overset grids to track dynamic flow features is described and the method is validated by tracking a moving shock and vortices shed behind a circular cylinder. Finally, Chimera gridding strategies were used to develop pressure coefficient contours for the FLAC wing for a Mach no. of 0.18 and Reynolds no. of 2.5 million.

Investigation of Vortex Flaps and Other Flow Control Devices on Generic High-Speed Civil Transport Planforms

NASA Technical Reports Server (NTRS)

Kjerstad, Kevin J.; Campbell, Bryan A.; Gile, Brenda E.; Kemmerly, Guy T.

1999-01-01

A parametric cranked delta planform study has been conducted in the Langley 14- by 22-Foot Subsonic Tunnel with the following objectives: (1) to evaluate the vortex flap design methodology for cranked delta wings, (2) to determine the influence of leading-edge sweep and the outboard wing on vortex flap effectiveness, (3) to evaluate novel flow control concepts, and (4) to validate unstructured grid Euler computer code predictions with modeled vortex and trailing-edge flaps. Two families of cranked delta planforms were investigated. One family had constant aspect ratio, while the other had a constant nondimensional semispan location of the leading-edge break. The inboard leading-edge sweep of the planforms was varied between 68 deg., 71 deg., and 74 deg., while outboard leading-edge sweep was varied between 48 deg. and 61 deg. Vortex flaps for the different planforms were designed by an analytical vortex flap design method. The results indicate that the effectiveness of the vortex flaps was only slightly influenced by the variations in the parametric planforms. The unstructured grid Euler computer code was successfully used to model the configurations with vortex flaps. The vortex trap concept was successfully demonstrated.

Unconventional missile concepts from consideration of varied mission requirements

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1984-01-01

Missile concepts for volumetric efficiency, minimum carriage constraints, and aerodynamic performance to achieve mission requirements. The mission requirements considered include air to surface roles such as defense suppression or antishipping where payload and range may have priority over high maneuver capability, and air to air and surface to air roles paying attention to good maneuvering capability. The concepts are intended to provide for ease of storage or carriage. The concepts include monoplanes with highly swept, thick delta wings, highly swept delta wings mounted either high or low on a semicircular body, some ring wing and semiring wing arrangements, parasol wing, and elliptical lifting bodies. The missile configurations indicate possible approaches toward resolving problems of carriage and storage while retaining good volumetric and aerodynamic efficiency. The configurations can accomplish a variety of possible missions with relatively simple vehicle shapes.

A low-speed wind tunnel study of vortex interaction control techniques on a chine-forebody/delta-wing configuration

NASA Technical Reports Server (NTRS)

Rao, Dhanvada M.; Bhat, M. K.

1992-01-01

A low speed wind tunnel evaluation was conducted of passive and active techniques proposed as a means to impede the interaction of forebody chine and delta wing vortices, when such interaction leads to undesirable aerodynamic characteristics particularly in the post stall regime. The passive method was based on physically disconnecting the chine/wing junction; the active technique employed deflection of inboard leading edge flaps. In either case, the intent was to forcibly shed the chine vortices before they encountered the downwash of wing vortices. Flow visualizations, wing pressures, and six component force/moment measurements confirmed the benefits of forced vortex de-coupling at post stall angles of attack and in sideslip, viz., alleviation of post stall zero beta asymmetry, lateral instability and twin tail buffet, with insignificant loss of maximum lift.

Exploratory investigation of the aerodynamic characteristics of a biwing vehicle at Mach 20.3

NASA Technical Reports Server (NTRS)

Bernot, P. T.

1984-01-01

Longitudinal and lateral-directional characteristics of a simple biwing configuration were determined over an angle-of-attack range from -3 deg to 50 deg. The body was comprised of a cylindrical section with an ogival forebody having an overall fineness ratio of 6.67. The delta wings had a 38.3 deg sweep angle and were geometrically similar in planform. The upper wing was located slightly forward relative to the lower wing. The model was tested in upright and inverted orientations including component buildups. This investigation was conducted in the 22-inch aerodynamics leg of the Langley Hypersonic Helium Tunnel Facility.

Influence of trap design on upwind flight behavior and capture of female grape berry moth (Lepidoptera: Tortricidae) with a kairomone lure.

PubMed

Cha, Dong H; Hesler, Stephen P; Linn, Charles E; Zhang, Aijun; Teal, Peter E A; Knight, Alan L; Roelofs, Wendell L; Loeb, Gregory M

2013-02-01

Oil-coated clear panel traps baited with a host plant-based kairomone lure have successfully been used for monitoring female grape berry moth, Paralobesia viteana (Clemens) (Lepidoptera: Tortricidae), but low capture rates as well as difficulty in servicing these traps makes them unsuitable for commercial use. We compared the performance of different trap designs in a flight tunnel and in a vineyard by using a 7-component synthetic kairomone blend, with a focus on trap visual cues. In flight tunnel experiments, a clear delta trap performed better than other traps. When we tested clear delta, green delta, or clear wing traps baited with a cut grape shoot, >50% of female grape berry moths made complete upwind flights. However, the clear delta trap was the only design that resulted in female moths entering the trap. Similar results were observed when females were tested with different traps (clear delta, green delta, white delta, clear wing, or green wing traps) baited with the kairomone lure. Adding a visual pattern that mimicked grape shoots to the outside surface of the clear delta trap resulted in 66% of the females that made upwind flights entering the trap. However, the positive effect of adding a visual pattern to the trap was not observed in a vineyard setting, where clear delta traps with or without a visual pattern caught similar numbers of females. Still, the number of male and female grape berry moths captured in clear delta traps with or without a visual pattern was not significantly different from the number of male and female grape berry moths captured in panel traps, suggesting that the use of these delta traps could be a less cumbersome alternative to oil-coated panel traps for monitoring female grape berry moth.

Aerodynamic characteristics of three slender sharp-edge 74 degrees swept wings at subsonic, transonic, and supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Davenport, E. E.

1974-01-01

Slender sharp-edge wings having leading-edge sweep angles of 74 deg have been studied at Mach numbers from 0.60 to 2.80, at angles of attack from about minus 4 deg to 22 deg, and at angles of sideslip from 0 deg to 5 deg. The wings had delta, arrow, and diamond planforms. The experimental tests were made in the Langley 8-foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel test section number 1. The theoretical predictions were made using the theories of NASA TN D-3767 and NASA TN D-6243. The results of the study indicated that the lift and drag characteristics as affected by planform and Mach number could be reasonably well predicted for the delta wing in the subsonic and transonic Mach number range. In the supersonic range, the delta and diamond wings were about equally good in the degree of agreement between experiment and theory. In making drag-due-to-lift predictions the vortex lift effects must be taken into account if reasonable results are to be obtained at moderate or high lift coefficients.

Delta wing vortex manipulation using pulsed and steady blowing during ramp pitching

NASA Technical Reports Server (NTRS)

Moreira, J.; Johari, H.

1995-01-01

The effectiveness of steady and pulsed blowing as a method of controlling delta wing vortices during ramp pitching has been investigated in flow visualization experiments conducted in a water tunnel. The recessed angled spanwise blowing technique was utilized for vortex manipulation. This technique was implemented on a beveled 60 delta wing using a pair of blowing ports located beneath the vortex core at 40% chord. The flow was injected primarily in the spanwise direction but was also composed of a component normal to the wing surface. The location of vortex burst was measured as a function of blowing intensity and pulsing frequency under static conditions, and the optimum blowing case was applied at three different wing pitching rates. Experimental results have shown that, when the burst location is upstream of the blowing port, pulsed blowing delays vortex breakdown in static and dynamic cases. Dynamic tests verified the existence of a hysteresis effect and demonstrated the improvements offered by pulsed blowing over both steady blowing and no-blowing scenarios. The application of blowing, at the optimum pulsing frequency, made the vortex breakdown location comparable in static and ramp pitch-up conditions.

Numerical studies of incompressible flow around delta and double-delta wings

NASA Technical Reports Server (NTRS)

Krause, E.; Liu, C. H.

1989-01-01

The subject has been jointly investigated at NASA Langley Research Center and the Aerodynamisches Institut of the RWTH Aachen over a substantial period. The aim of this investigation has been to develop numerical integration procedures for the Navier-Stokes equations - particularly for incompressible three-dimensional viscous flows about simple and double delta wings - and to study the low speed flow behavior, with its complex vortex structures on the leeward side of the wing. The low speed flight regime poses unusual problems because high incidence flight conditions may, for example, encounter symmetric and asymmetric vortex breakdown. Because of the many difficulties to be expected in solving the problem, it was divided into two - analysis of the flow without vortex breakdown and analysis of the breakdown of isolated vortices. The major results obtained so far on the two topics are briefly described.

Experimental and numerical analysis of the wing rock characteristics of a 'wing-body-tail' configuration

NASA Technical Reports Server (NTRS)

Suarez, Carlos J.; Smith, Brooke C.; Malcolm, Gerald N.

1993-01-01

Free-to-roll wind tunnel tests were conducted and a computer simulation exercise was performed in an effort to investigate in detail the mechanism of wing rock on a configuration that consisted of a highly-slender forebody and a 78 deg swept delta wing. In the wind tunnel test, the roll angle and wing surface pressures were measured during the wing rock motion. A limit cycle oscillation was observed for angles of attack between 22 deg and 30 deg. In general, the wind tunnel test confirmed that the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. The variation of roll acceleration (determined from the second derivative of the roll angle time history) with roll angle clearly showed the energy balance necessary to sustain the limit cycle oscillation. Pressure measurements on the wing revealed the hysteresis of the wing rock process. First, second and nth order models for the aerodynamic damping were developed and examined with a one degree of freedom computer simulation. Very good agreement with the observed behavior from the wind tunnel was obtained.

Analytical and Experimental Investigations of Delta Wings in Incompressible Flow

DTIC Science & Technology

1976-08-01

posi- tion unless so designated by other official documents. Rep roduct ion Reproduction in whole or in part is permitted for any purpose of the...Trailing Edge of Free-Wake Model 56 19 Polar Coordinate System 5i 20 Free-Wake Geometry Prediction 5 21 Finite-Size Core 58 22 Vortex Core Position of Smith...k=0.4 70 34 Details of a Helical Type Burst 71 35 Vortex Burst-Steady Flow 72 36 Location of Pressure Ports 73 37 Pressure Destribution on a Delta

Space shuttle: Static aerodynamic characteristics and control effectiveness of two delta wing orbiter configurations (M equals 0.6 to 4.96)

NASA Technical Reports Server (NTRS)

Ellis, R. R.; Buchholz, R. E.; Moore, J. A.

1972-01-01

Two 0.00325-scale models of a space shuttle orbiter were tested in trisonic wind tunnel to obtain force, static stability, and control effectiveness data by six component internal strain gauge balance. Two separate configurations were tested; however, the fuselage and basic wing were of one-piece construction. The configurations were varied by replacing the straight wing tip extensions with upswept wing tips. Directional stability was provided for one configuration by a centerline vertical tail. Due to the one-piece body/wing construction, no body-alone data were obtained. The effect of tip fins and vertical tail size were, however, investigated. Both configurations were tested over a Mach range of 0.6 to 4.96 with data taken at angles of attack from minus 4 deg to 60 deg and at angles of sideslip from minus 4 deg to 10 deg.

Forebody vortex control for suppressing wing rock on a highly-swept wing configuration

NASA Technical Reports Server (NTRS)

Suarez, Carlos J.; Kramer, Brian R.; Ayers, Bert; Malcolm, Gerald N.

1992-01-01

Free-to-roll tests were conducted in a wind tunnel with a configuration that consisted of a highly-slender forebody and a 78 deg swept delta wing. A limit cycle oscillation was observed for angles of attack between 22 and 30 deg. In general, the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. Various blowing techniques were evaluated as means of wing rock suppression. Blowing tangentially aft from leeward side nozzles near the forebody tip can damp the roll motion at low blowing rates and stop it completely at higher blowing rates. At the high rates, significant vortex asymmetries are created, causing the model to stop at a non-zero roll angle. Forward blowing and alternating right/left pulsed blowing appear to be more efficient techniques for suppressing wing rock. The oscillations can be damped almost completely at lower blowing coefficients, and, apparently, no major vortex asymmetries are induced. Good agreement is observed between this study and previous water tunnel tests on the same configuration.

76 FR 78805 - Airworthiness Directives; Rolls-Royce plc (RR) RB211-Trent 800 Series Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2011-12-20

... rejected fuel tubes need to be replaced to avoid fuel leaks. Delta indicated that the On- wing Inspection... Revision Number Three commenters, American Airlines (American), Delta Airlines (Delta), and the Boeing... Revisions of the Service Bulletin One commenter, Delta, asked that engines previously inspected per Revision...

An experimental study of pressures on 60 deg Delta wings with leading edge vortex flaps

NASA Technical Reports Server (NTRS)

Marchman, J. F., III; Terry, J. E.; Donatelli, D. A.

1983-01-01

An experimental study was conducted in the Virginia Tech Stability Wind Tunnel to determine surface pressures over a 60 deg sweep delta wing with three vortex flap designs. Extensive pressure data was collected to provide a base data set for comparison with computational design codes and to allow a better understanding of the flow over vortex flaps. The results indicated that vortex flaps can be designed which will contain the leading edge vortex with no spillage onto the wing upper surface. However, the tests also showed that flaps designed without accounting for flap thickness will not be optimum and the result can be oversized flaps, early flap vortex reattachment and a second separation and vortex at the wing/flap hinge line.

Some aspects of hybrid-zeppelins. [optimization of delta wings for airships

NASA Technical Reports Server (NTRS)

Mackrodt, P. A.

1975-01-01

To increase an airship's maneuverability and payload capacity as well as to save bouyant gas it is proposed to outfit it with a slender delta-wing, which carries about one half of the total take-off weight of the vehicle. An optimization calculation based on the data of LZ 129 (the last airship, which saw passenger-service) leads to a Hybrid-Zeppelin with a wing of aspect-ratio 1.5 and 105 m span. The vehicle carries a payload of 40% of it's total take-off weight and consumes 0.8 t fuel per ton payload over a distance of 10000 km.

Reynolds Number, Compressibility, and Leading-Edge Bluntness Effects on Delta-Wing Aerodynamics

NASA Technical Reports Server (NTRS)

Luckring, James M.

2004-01-01

An overview of Reynolds number, compressibility, and leading edge bluntness effects is presented for a 65 degree delta wing. The results of this study address both attached and vortex-flow aerodynamics and are based upon a unique data set obtained in the NASA-Langley National Transonic Facility (NTF) for i) Reynolds numbers ranging from conventional wind-tunnel to flight values, ii) Mach numbers ranging from subsonic to transonic speeds, and iii) leading-edge bluntness values that span practical slender wing applications. The data were obtained so as to isolate the subject effects and they present many challenges for Computational Fluid Dynamics (CFD) studies.

Effectiveness of Flow Control for Alleviation of Twin-Tail Buffet

NASA Technical Reports Server (NTRS)

Sheta, Essam F.; Kandil, Osama A.; Yang, Zhi

1998-01-01

Effectiveness of active flow control for twin- tail buffet alleviation is investigated. Tangen- tial leading-edge blowing (TLEB) and flow suction along the vortex cores (FSVC) of the lead- ing edges of the delta wing are used to delay the vortex breakdown flow upstream of the twin tail. The combined effect of the TLEB and FSVC is also investigated. A parametric study of the effects of the spanwise position of the suction tubes and volumetric suction flow rate on the twin-tail buffet response are also investigated. The TLEB moves the path of leading-edge vortices laterally towards the twin tail, which increases the aero- dynamic damping on the tails. The FSVC effectively delays the breakdown location at high angles of attack. The computational model consists of a sharp-edged delta wing of aspect ratio one and swept-back flexible twin tail with taper ratio of 0.23. This complex multidisciplinary problem is solved sequentially using three sets of equations for the fluid flow, aeroelastic response and grid deformation, on a dynamic multi-block grid structure. The computational model is pitched at 30 deg. angle of attack. The freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. The model is investigated for the inboard position of the twin tails, which corresponds to a separation distance between the twin tails of 33% of the wing span.

Pressure-Sensitive Paint Investigation of Double-Delta Wing Vortex Flow Manipulation

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Gonzalez, Hugo A.

2004-01-01

A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76o/40o double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M =0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

Pressure-Sensitive Paint Investigation of Double-Delta Wing Vortex Flow Manipulation

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Gonzalez, Hugo A.

2005-01-01

A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to quantify the effect of wing fillets on the global vortex-induced surface static pressure field about a sharp leading-edge 76 deg/40 deg double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 30 degrees using an in-situ method featuring the simultaneous acquisition of electronically-scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M = 0.50 to 0.85 but increased to several percent at M = 0.95 and 1.20. The PSP pressure distributions and pseudo-colored planform view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having a parabolic or diamond planform situated at the strake-wing intersection were designed to manipulate the vortical flows by, respectively, removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

On the nonlinear aerodynamic and stability characteristics of a generic chine-forebody slender-wing fighter configuration

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Brandon, Jay M.

1987-01-01

An exploratory investigation was conducted of the nonlinear aerodynamic and stability characteristics of a tailless generic fighter configuration featuring a chine-shaped forebody coupled to a slender cropped delta wing in the NASA Langley Research Center's 12-Foot Low-Speed Wind Tunnel. Forebody and wing vortex flow mechanisms were identified through off-body flow visualizations to explain the trends in the longitudinal and lateral-directional characteristics at extreme attitudes (angles of attack and sideslip). The interactions of the vortical motions with centerline and wing-mounted vertical tail surfaces were studied and the flow phenomena were correlated with the configuration forces and moments. Single degree of freedom, free-to-roll tests were used to study the wing rock susceptibility of the generic fighter model. Modifications to the nose region of the chine forebody were examined and fluid mechanisms were established to account for their ineffectiveness in modulating the highly interactive forebody and wing vortex systems.

Exploratory Wind-Tunnel Investigation to Determine the Lift Effects of Blowing over Flaps from Nacelles Mounted Above the Wing

NASA Technical Reports Server (NTRS)

Riebe, John M; Davenport, Edwin E

1958-01-01

An exploratory wind-tunnel investigation has been made to determine the lift effects of blowing from nacelles over the upper surface of flaps on a model having a delta wing of aspect ratio 3. Several flap conditions were examined. High-pressure air was blown from an external-pipe arrangement supported above the wing to simulate jet-engine exhaust. The jet momentum- coefficient range was from 0 to 3.0 and the model angle of attack was 0 deg. The results of this limited investigation show that values of jet circulation lift coefficient larger than the Jet reaction were produced with blowing over flaps from nacelles mounted above the wing. 'I!heuse of double slotted flaps with the gap unsealed between the flaps and wing had a large detrimental effect on the lift capabilities. With these gaps sealed, larger lift coefficients were obtained when fantails were added to the nacelles. The longitudinal trim problems created by large diving moments were similar to those encountered with other jet-augmented-flap systems

78 FR 24785 - Public Hearing

Federal Register 2010, 2011, 2012, 2013, 2014

2013-04-26

... hearing will be conducted at the Pennsylvania State Capitol, Room 8E-B, East Wing, Commonwealth Avenue....283 mgd (peak day). 9. Project Sponsor: Delta Borough Municipal Authority. Project Facility: Delta... to 0.073 mgd (30-day average) from Well 5. 10. Project Sponsor: Delta Borough Municipal Authority...

Motion simulator study of longitudinal stability requirements for large delta wing transport airplanes during approach and landing with stability augmentation systems failed

NASA Technical Reports Server (NTRS)

Snyder, C. T.; Fry, E. B.; Drinkwater, F. J., III; Forrest, R. D.; Scott, B. C.; Benefield, T. D.

1972-01-01

A ground-based simulator investigation was conducted in preparation for and correlation with an-flight simulator program. The objective of these studies was to define minimum acceptable levels of static longitudinal stability for landing approach following stability augmentation systems failures. The airworthiness authorities are presently attempting to establish the requirements for civil transports with only the backup flight control system operating. Using a baseline configuration representative of a large delta wing transport, 20 different configurations, many representing negative static margins, were assessed by three research test pilots in 33 hours of piloted operation. Verification of the baseline model to be used in the TIFS experiment was provided by computed and piloted comparisons with a well-validated reference airplane simulation. Pilot comments and ratings are included, as well as preliminary tracking performance and workload data.

Survey of research on unsteady aerodynamic loading of delta wings

NASA Technical Reports Server (NTRS)

Ashley, H.; Vaneck, T.; Katz, J.; Jarrah, M. A.

1991-01-01

For aeronautical applications, there has been recent interest in accurately determining the aerodynamic forces and moments experienced by low-aspect-ratio wings performing transient maneuvers which go to angles of attack as high as 90 deg. Focusing on the delta planform with sharp leading edges, the paper surveys experimental and theoretical investigations dealing with the associated unsteady flow phenomena. For maximum angles above a value between 30 and 40 deg, flow details and airloads are dominated by hysteresis in the 'bursting' instability of intense vortices which emanate from the leading edge. As examples of relevant test results, force and moment histories are presented for a model series with aspect ratios 1, 1.5 and 2. Influences of key parameters are discussed, notably those which measure unsteadiness. Comparisons are given with two theories: a paneling approximation that cannot capture bursting but clarifies other unsteady influences, and a simplified estimation scheme which uses measured bursting data.

Space shuttle: Stability and control effectiveness of the MDAC parametric delta canard booster at Mach 0.38. Volume 1: Canard parametric variations

NASA Technical Reports Server (NTRS)

Bradley, D.; Buchholz, R. E.

1971-01-01

A 0.015 scale model of a modified version of the MDAC space shuttle booster was tested in the Naval Ship Research and Development Center 7 x 10 foot transonic wind tunnel, to obtain force, static stability, and control effectiveness data. Data were obtained for a cruise Mach Number of 0.38, altitude of 10,000 ft, and Reynolds Number per foot of approximately 2 x one million. The model was tested through an angle of attack range of -4 deg to 15 deg at zero degree angle of sideslip, and at an angle of sideslip range of -6 deg to 6 deg at fixed angles of attack of 0 deg, 6 deg, and 15 deg. Other test variables were elevon deflections, canard deflections, aileron deflections, rudder deflections, wing dihedral angle, canard incidence angle, wing incidence angle, canard position, wing position, wing and canard control flap size and dorsal fin size.

Reynolds number effects on the aerodynamic characteristics of irregular planform wings at Mach number 0.3. [in the Ames 12 ft pressure wind tunnel

NASA Technical Reports Server (NTRS)

Kruse, R. L.; Lovette, G. H.; Spencer, B., Jr.

1977-01-01

The subsonic aerodynamic characteristics of a series of irregular planform wings were studied in wind tunnel tests conducted at M = 0.3 over a range of Reynolds numbers from 1.6 million to 26 million/m. The five basic wing planforms varied from a trapezoidal to a delta shape. Leading edge extensions, added to the basic shape, varied in approximately 5 deg increments from the wing leading edge sweep-back angle to a maximum 80 deg. Most of the tests were conducted using an NACA 0008 airfoil section with grit boundary layer trips. Tests were also conducted using an NACA 0012 airfoil section and an 8% thick wedge. In addition, the effect of free transition (no grit) was investigated. A body was used on all models.

78 FR 8131 - Federal Home Loan Bank Members Selected for Community Support Review

Federal Register 2010, 2011, 2012, 2013, 2014

2013-02-05

... Union Chelsea Massachusetts. Pilgrim Bank Cohasset Massachusetts. Delta-Wye Federal Credit Union... Minnesota. The First State Bank of Red Wing......... Red Wing Minnesota. Eastwood Bank Rochester Minnesota...

75 FR 16463 - Federal Home Loan Bank Members Selected for Community Support Review

Federal Register 2010, 2011, 2012, 2013, 2014

2010-04-01

... Chelsea Massachusetts. Pilgrim Bank Cohasset Massachusetts. Delta-Wye Federal Credit Union Dorchester... Community Bank, National Preston Minnesota. Association. First State Bank of Red Wing (The)....... Red Wing...

Reading Ombrone river delta evolution through beach ridges morphology

NASA Astrophysics Data System (ADS)

Mammi, Irene; Piccardi, Marco; Pranzini, Enzo; Rossi, Lorenzo

2017-04-01

The present study focuses on the evolution of the Ombrone River delta (Southern Tuscany, Italy) in the last five centuries, when fluvial sediment input was huge also as a consequence of the deforestation performed on the watershed. The aim of this study is to find a correlation between river input and beach ridges morphology and to explain the different distribution of wetlands and sand deposits on the two sides of the delta. Visible, NIR and TIR satellite images were processed to retrieve soil wetness associated to sand ridges and interdune silty deposits. High resolution LiDAR data were analysed using vegetation filter and GIS enhancement algorithms in order to highlight small morphological variations, especially in areas closer to the river where agriculture has almost deleted these morphologies. A topographic survey and a very high resolution 3D model obtained from a set of images acquired by an Unmanned Aerial Vehicle (UAV) were carried out in selected sites, both to calibrate satellite LiDAR 3D data, and to map low relief areas. Historical maps, aerial photography and written documents were analysed for dating ancient shorelines associated to specific beach ridges. Thus allowing the reconstruction of erosive and accretive phases of the delta. Seventy beach ridges were identified on the two wings of the delta. On the longer down-drift side (Northern wing) beach ridges are more spaced at the apex and gradually converge to the extremity, where the Bruna River runs and delimits the sub aerial depositional area of the Ombrone River. On the shorter up-drift lobe (Southern wing), beach ridges are closer, but run almost parallel each other. In this case, a rocky headland called Collelungo promontory closes and cuts the beach ridges sequence but shallow water depth allows sediment by pass. One kilometre to the south a more pronounced promontory encloses a small pocket beach (Cala di Forno) and identifies the limit of the subaerial depositionary area. Beach ridges heights were analysed through LiDAR data and some of them were found higher than average. Conceptual models in literature allowed us to explain higher beach ridges as periods of stability or a very initial erosion stage interesting the beach. The high resolution DTM produced from LiDAR and UAV data permitted a better reconstruction of the last five centuries of delta evolution and to characterize the difference of beach ridges morphology of the up-drift and the down-drift sides of the delta. Within this framework the presence of interdune swales in the down-drift side has been explained.

Wind Tunnel Application of a Pressure-Sensitive Paint Technique to a Double Delta Wing Model at Subsonic and Transonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Gonzalez, Hugo A.

2006-01-01

A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to study the effect of wing fillets on the global vortex induced surface static pressure field about a sharp leading-edge 76 deg./40 deg. double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M(sub infinity) = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an insitu method featuring the simultaneous acquisition of electronically scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M(sub infinity) = 0.50 to 0.85 but increased to several percent at M(sub infinity) =0.95 and 1.20. The PSP pressure distributions and pseudo-colored, planform-view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having parabolic or diamond planforms situated at the strake-wing intersection were respectively designed to manipulate the vortical flows by removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.

14 CFR Appendix A to Part 23 - Simplified Design Load Criteria

Code of Federal Regulations, 2011 CFR

2011-01-01

... quarter-chord), delta planforms, or slatted lifting surfaces; or (5) Winglets or other wing tip devices... single engine excluding turbine powerplants; (2) A main wing located closer to the airplane's center of gravity than to the aft, fuselage-mounted, empennage; (3) A main wing that contains a quarter-chord sweep...

77 FR 64696 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-10-23

... installing an unmodified wing panel on that airplane. Delta justified its request by stating that the lack of... following a ground fire incident and reports of flammable fluid leaks from the wing leading edge area onto... area of the wing. We are issuing this AD to prevent flammable fluid from leaking onto the engine...

14 CFR Appendix A to Part 23 - Simplified Design Load Criteria

Code of Federal Regulations, 2010 CFR

2010-01-01

... quarter-chord), delta planforms, or slatted lifting surfaces; or (5) Winglets or other wing tip devices... single engine excluding turbine powerplants; (2) A main wing located closer to the airplane's center of gravity than to the aft, fuselage-mounted, empennage; (3) A main wing that contains a quarter-chord sweep...

Control of vortex on a non-slender delta wing by a nanosecond pulse surface dielectric barrier discharge

NASA Astrophysics Data System (ADS)

Zhao, Guang-yin; Li, Ying-hong; Liang, Hua; Han, Meng-hu; Hua, Wei-zhuo

2015-01-01

Wind tunnel experiments are conducted for improving the aerodynamic performance of delta wing using a leading-edge pulsed nanosecond dielectric barrier discharge (NS-DBD). The whole effects of pulsed NS-DBD on the aerodynamic performance of the delta wing are studied by balanced force measurements. Pressure measurements and particle image velocimetry (PIV) measurements are conducted to investigate the formation of leading-edge vortices affected by the pulsed NS-DBD, compared to completely stalled flow without actuation. Various pulsed actuation frequencies of the plasma actuator are examined with the freestream velocity up to 50 m/s. Stall has been delayed substantially and significant shifts in the aerodynamic forces can be achieved at the post-stall regions when the actuator works at the optimum reduced frequency of F + = 2. The upper surface pressure measurements show that the largest change of static pressure occurs at the forward part of the wing at the stall region. The time-averaged flow pattern obtained from the PIV measurement shows that flow reattachment is promoted with excitation, and a vortex flow pattern develops. The time-averaged locations of the secondary separation line and the center of the vortical region both move outboard with excitation.

Computational Aerodynamics of Shuttle Orbiter Damage Scenarios in Support of the Columbia Accident Investigation

NASA Technical Reports Server (NTRS)

Bibb, Karen L.; Prabhu, Ramadas K.

2004-01-01

In support of the Columbia Accident Investigation, inviscid computations of the aerodynamic characteristics for various Shuttle Orbiter damage scenarios were performed using the FELISA unstructured CFD solver. Computed delta aerodynamics were compared with the reconstructed delta aerodynamics in order to postulate a progression of damage through the flight trajectory. By performing computations at hypervelocity flight and CF4 tunnel conditions, a bridge was provided between wind tunnel testing in Langley's 20-Inch CF4 facility and the flight environment experienced by Columbia during re-entry. The rapid modeling capability of the unstructured methodology allowed the computational effort to keep pace with the wind tunnel and, at times, guide the wind tunnel efforts. These computations provided a detailed view of the flowfield characteristics and the contribution of orbiter components (such as the vertical tail and wing) to aerodynamic forces and moments that were unavailable from wind tunnel testing. The damage scenarios are grouped into three categories. Initially, single and multiple missing full RCC panels were analyzed to determine the effect of damage location and magnitude on the aerodynamics. Next is a series of cases with progressive damage, increasing in severity, in the region of RCC panel 9. The final group is a set of wing leading edge and windward surface deformations that model possible structural deformation of the wing skin due to internal heating of the wing structure. By matching the aerodynamics from selected damage scenarios to the reconstructed flight aerodynamics, a progression of damage that is consistent with the flight data, debris forensics, and wind tunnel data is postulated.

Myoblast cytonemes mediate Wg signaling from the wing imaginal disc and Delta-Notch signaling to the air sac primordium.

PubMed

Huang, Hai; Kornberg, Thomas B

2015-05-07

The flight muscles, dorsal air sacs, wing blades, and thoracic cuticle of the Drosophila adult function in concert, and their progenitor cells develop together in the wing imaginal disc. The wing disc orchestrates dorsal air sac development by producing decapentaplegic and fibroblast growth factor that travel via specific cytonemes in order to signal to the air sac primordium (ASP). Here, we report that cytonemes also link flight muscle progenitors (myoblasts) to disc cells and to the ASP, enabling myoblasts to relay signaling between the disc and the ASP. Frizzled (Fz)-containing myoblast cytonemes take up Wingless (Wg) from the disc, and Delta (Dl)-containing myoblast cytonemes contribute to Notch activation in the ASP. Wg signaling negatively regulates Dl expression in the myoblasts. These results reveal an essential role for cytonemes in Wg and Notch signaling and for a signal relay system in the myoblasts.

Fourier functional analysis for unsteady aerodynamic modeling

NASA Technical Reports Server (NTRS)

Lan, C. Edward; Chin, Suei

1991-01-01

A method based on Fourier analysis is developed to analyze the force and moment data obtained in large amplitude forced oscillation tests at high angles of attack. The aerodynamic models for normal force, lift, drag, and pitching moment coefficients are built up from a set of aerodynamic responses to harmonic motions at different frequencies. Based on the aerodynamic models of harmonic data, the indicial responses are formed. The final expressions for the models involve time integrals of the indicial type advocated by Tobak and Schiff. Results from linear two- and three-dimensional unsteady aerodynamic theories as well as test data for a 70-degree delta wing are used to verify the models. It is shown that the present modeling method is accurate in producing the aerodynamic responses to harmonic motions and the ramp type motions. The model also produces correct trend for a 70-degree delta wing in harmonic motion with different mean angles-of-attack. However, the current model cannot be used to extrapolate data to higher angles-of-attack than that of the harmonic motions which form the aerodynamic model. For linear ramp motions, a special method is used to calculate the corresponding frequency and phase angle at a given time. The calculated results from modeling show a higher lift peak for linear ramp motion than for harmonic ramp motion. The current model also shows reasonably good results for the lift responses at different angles of attack.

77 FR 12444 - Airworthiness Directives; General Electric Company (GE) Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-03-01

... support the NPRM as written. Request To Correct Part Number Commenters GE and Delta Airlines (Delta... packing configuration either on-wing or in the shop and to determine if further actions are necessary. We.... Request To Modify Compliance Time Commenters GE and Delta requested that we change the compliance time of...

77 FR 32433 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-06-01

... front spar of the wing through the slat track housing. This action revises that NPRM by adding...-on actions. Requests To Clarify Reporting Results Boeing, Delta Air Lines (Delta), and FedEx... request is not shown in paragraph (h) of the NPRM (75 FR 31327, June 3, 2010). Delta stated that reporting...

Integrated Experimental and Numerical Research on the Aerodynamics of Unsteady Moving Aircraft

DTIC Science & Technology

2007-06-01

blended wing body configuration were tested in different modes of oscillatory motions roll, pitch and yaw as well as delta wing geometries like X-31...airplane configurations (e.g. wide body, green aircraft, blended wing body) the approach up to now using semi-empirical methods as standard...cross section wing. In order to evaluate the influence of individual components of the tested airplane configuration, such as winglets , vertical or

Convair XF-102 Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1953-08-21

A .10-scale model of Convair’s XF-102 in the 8- by 6-Foot Supersonic Wind Tunnel at the National Advisory Committee for Aeronautics (NACA) Lewis Flight Propulsion Laboratory for jet exit studies. The XF-102 was a prototype of the F-102 Delta Dagger. The F-102 served as an interceptor against long range bombers from the Soviet Union. The aircraft was powered by a Pratt and Whitney J57 turbojet. The first prototype crashed two weeks after is first flight on October 24, 1953, just months after this photograph. Engineers then incorporated the fixed-wing design to reduce drag at supersonic speeds. The production model F-102 became the first delta-wing supersonic aircraft in operation. The 8- by 6-Foot Supersonic Wind Tunnel is used to study propulsion systems, including inlets and exit nozzles, combustion fuel injectors, flame holders, exit nozzles, and controls on ramjet and turbojet engines. Flexible sidewalls alter the tunnel’s nozzle shape to vary the Mach number during operation. A seven-stage axial compressor, driven by three electric motors that yield a total of 87,000 horsepower, generates air speeds from Mach 0.36 to 2.0.

Recent transonic unsteady pressure measurements at the NASA Langley Research Center

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Ricketts, R. H.; Hess, R. W.

1985-01-01

Four semispan wing model configurations were studied in the Transonic Dynamics Tunnel (TDT). The first model had a clipped delta planform with a circular arc airfoil, the second model had a high aspect ratio planform with a supercritical airfoil, the third model has a rectangular planform with a supercritical airfoil and the fourth model had a high aspect ratio planform with a supercritical airfoil. To generate unsteady flow, the first and third models were equipped with pitch oscillation mechanisms and the first, second and fourth models were equipped with control surface oscillation mechanisms. The fourth model was similar in planform and airfoil shape to the second model, but it is the only one of the four models that has an elastic wing structure. The unsteady pressure studies of the four models are described and some typical results for each model are presented. Comparison of selected experimental data with analytical results also are included.

Leeward flow over delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Szodruch, J. G.

1980-01-01

A survey was made of the parameters affecting the development of the leeward symmetric separated flow over slender delta wings immersed in a supersonic stream. The parameters included Mach number, Reynolds number, angle of attack, leading-edge sweep angle, and body cross-sectional shape, such that subsonic and supersonic leading-edge flows are encountered. It was seen that the boundaries between the various flow regimes existing about the leeward surface may conveniently be represented on a diagram with the components of angle of attack and Mach number normal to the leading edge as governing parameters.

Computer program for calculating supersonic flow on the windward side conical delta wings by the method of lines

NASA Technical Reports Server (NTRS)

Klunker, E. B.; South, J. C., Jr.; Davis, R. M.

1972-01-01

A user's manual is presented for a program that calculates the supersonic flow on the windward side of conical delta wings with shock attached at the sharp leading edge by the method of lines. The program also has a limited capability for computing the flow about circular and elliptic cones at incidence. It provides information including the shock shape, flow field, isentropic surface-flow properties, and force coefficients. A description of the program operation, a sample computation, and a FORTRAN 4 program listing are included.

Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading edge vortex separation.

Reynolds Number and Leading-Edge Bluntness Effects on a 65 Deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2002-01-01

A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M = 0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading- edge vortex separation.

Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2002-01-01

A 65 degree delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at subsonic speeds (M = 0.4) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

Transonic Reynolds Number and Leading-Edge Bluntness Effects on a 65 deg Delta Wing

NASA Technical Reports Server (NTRS)

Luckring, J. M.

2003-01-01

A 65 deg delta wing has been tested in the National Transonic Facility (NTF) at mean aerodynamic chord Reynolds numbers from 6 million to 120 million at subsonic and transonic speeds. The configuration incorporated a systematic variation of the leading edge bluntness. The analysis for this paper is focused on the Reynolds number and bluntness effects at transonic speeds (M=0.85) from this data set. The results show significant effects of both these parameters on the onset and progression of leading-edge vortex separation.

Aerodynamic studies of delta-wing shuttle orbiters. Part 1: Low speed

NASA Technical Reports Server (NTRS)

Freeman, D. C., Jr.; Ellison, J. C.

1972-01-01

Numerous wind tunnel tests conducted on the evolving delta-wing orbiters have generated a fairly large aerodynamic data base over the entire entry operation range of these vehicles. A limited assessment is made of some of the aerodynamics of the current HO type orbiters, and several specific problem areas selected from the broad data base are discussed. These include, from a subsonic viewpoint, discussions of trim drag effect; effects of the installation of main rocket engine nozzles, OMS and RCS packages, Reynolds number effects, lateral-directional stability characteristics, and landing characteristics.

Vortical flow management for improved configuration aerodynamics: Recent experiences

NASA Technical Reports Server (NTRS)

Rao, D. M.

1983-01-01

Recent progress in vortex-control applications for alleviating the adverse consequences of three dimensional separation and vortical interactions on slender body/swept wing configurations is reported. Examples include helical separation trip to alleviate the side force due to forebody vortex asymmetry; hinged strakes to avoid vortex breakdown effects; compartmentation of swept leading edge separation to delay the pitch-up instability; under wing vortex trip and vortex trip and vortex flaps for drag reduction at high lift; and an apex-flap trimmer to fully utilize the lift capability of trailing-edge flaps for take off and landing of delta wings. Experimental results on generic wind-tunnel models are presented to illustrate the vortex-management concepts involved and to indicate their potential for enhancing the subsonic aerodynamics of supersonic-cruise type vehicles.

An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

NASA Technical Reports Server (NTRS)

Ghaffari, F.; Chaturvedi, S. K.

1984-01-01

An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

An investigation of empennage buffeting

NASA Technical Reports Server (NTRS)

Lan, C. E.; Lee, I. G.

1986-01-01

Progress in the investigation of empennage buffeting in reviewed. In summary, the following tasks were accomplished: relevant literatures was reviewed; equations for calculating structural response were formulated; root-mean-square values of root bending moment for a 65-degree rigid delta wing were calculated and compared with data; and a water-tunnel test program for an F-18 model was completed.

A numerical study of the effect of geometry variation, turbulence models, and dissipation on the flow past control surfaces

NASA Technical Reports Server (NTRS)

Chaussee, Denny S.

1993-01-01

The steady 3D viscous flow past the ONERA M6 wing and a slender delta wing-body with trailing edge control surfaces has been computed. A cell-centered finite-volume Navier-Stokes patched zonal method has been used for the numerical simulation. Both diagonalized and LUSGS schemes have been implemented. Besides the standard nonplanar zonal interfacing techniques, a new virtual zone capability has been employed. For code validation, the transonic flow past the ONERA M5 wing is calculated for angles-of-attack of 3.06 deg and 5.06 deg and compared with the available experiments. The wing-body computational results are compared with experimental data for both trailing-edge flaps deflected. The experimental flow conditions are M subinfinity = 0.4, a turbulent Reynolds number of 5.41 million based on a mean aerodynamic chord of 25.959 inches, adiabatic wall, and angles-of-attack varying from 0 deg to 23.85 deg. The computational results are presented for the 23.85 deg angle-of-attack case. The effects of the base flow due to a model sting, the varying second and fourth order numerical dissipation, and the turbulence model are all considered.

Pressure investigation of NASA leading edge vortex flaps on a 60 deg Delta wing

NASA Technical Reports Server (NTRS)

Marchman, J. F., III; Donatelli, D. A.; Terry, J. E.

1983-01-01

Pressure distributions on a 60 deg Delta Wing with NASA designed leading edge vortex flaps (LEVF) were found in order to provide more pressure data for LEVF and to help verify NASA computer codes used in designing these flaps. These flaps were intended to be optimized designs based on these computer codes. However, the pressure distributions show that the flaps wre not optimum for the size and deflection specified. A second drag-producing vortex forming over the wing indicated that the flap was too large for the specified deflection. Also, it became apparent that flap thickness has a possible effect on the reattachment location of the vortex. Research is continuing to determine proper flap size and deflection relationships that provide well-behaved flowfields and acceptable hinge-moment characteristics.

Comparisons of two-dimensional shock-expansion theory with experimental aerodynamic data for delta-planform wings at high supersonic speeds

NASA Technical Reports Server (NTRS)

Jernell, L. S.

1974-01-01

An investigation has been conducted to explore the potential for optimizing airfoil shape at high supersonic speeds by utilizing the two-dimensional shock-expansion method. Theoretical and experimental force and moment coefficients are compared for four delta-planform semispan wings having a leading-edge sweep angle of 65 deg and incorporating modified diamond airfoils with a thickness-chord ratio of 0.06. The wings differ only in airfoil maximum-thickness position and camber. The experimental data are obtained at Mach numbers of 3.95 and 4.63 and at a Reynolds number of 9.84 million per meter. A relatively simple method is developed for predicting, in terms of lift-drag ratio, the optimum modified diamond airfoil at high supersonic and hypersonic speeds.

Unsteady transonic flow analysis for low aspect ratio, pointed wings.

NASA Technical Reports Server (NTRS)

Kimble, K. R.; Ruo, S. Y.; Wu, J. M.; Liu, D. Y.

1973-01-01

Oswatitsch and Keune's parabolic method for steady transonic flow is applied and extended to thin slender wings oscillating in the sonic flow field. The parabolic constant for the wing was determined from the equivalent body of revolution. Laplace transform methods were used to derive the asymptotic equations for pressure coefficient, and the Adams-Sears iterative procedure was employed to solve the equations. A computer program was developed to find the pressure distributions, generalized force coefficients, and stability derivatives for delta, convex, and concave wing planforms.

Large-Amplitude, High-Rate Roll Oscillations of a 65 deg Delta Wing at High Incidence

NASA Technical Reports Server (NTRS)

Chaderjian, Neal M.; Schiff, Lewis B.

2000-01-01

The IAR/WL 65 deg delta wing experimental results provide both detail pressure measurements and a wide range of flow conditions covering from simple attached flow, through fully developed vortex and vortex burst flow, up to fully-stalled flow at very high incidence. Thus, the Computational Unsteady Aerodynamics researchers can use it at different level of validating the corresponding code. In this section a range of CFD results are provided for the 65 deg delta wing at selected flow conditions. The time-dependent, three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate the unsteady vertical flow. Two sting angles and two large- amplitude, high-rate, forced-roll motions and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are compared with experimental pressures, forces, moments and roll angle time history. In addition, surface and off-surface flow particle streaks are also presented.

Effect of Configuration Pitching Motion on Twin Tail Buffet Response

NASA Technical Reports Server (NTRS)

Sheta, Essam F.; Kandil, Osama A.

1998-01-01

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.

Exploring Flight Research with Experimental Gliders

NASA Technical Reports Server (NTRS)

1999-01-01

A look at the research aircraft flown by NASA and its predecessor, the National Advisory Committee for Aeronautics (NACA), since the 1940's reveals an evolution of wing designs. In fact, each of the first series of NACA experimental research aircraft ("X-planes") used different wing and tail configurations to tackle the problems of supersonic flight, These early jet aircraft had straight wings (X-1), wings that angled (swept) toward the tail (X-2), triangular (delta) wings (XF-92), and wings that could be moved in flight to change the angle of backward sweep (X-5). Each design added to our knowledge of high-speed flight.

Effect of Dynamic Rolling Oscillations on Twin Tail Buffet Response

NASA Technical Reports Server (NTRS)

Sheta, Essam F.; Kandil, Osama A.

1999-01-01

The effect of dynamic rolling oscillations of delta-wing/twin-tail configuration on twin-tail buffet response is investigated. The computational model consists of a sharp-edged delta wing of aspect ratio one and swept-back flexible twin tail with taper ratio of 0.23. The configuration model is statically pitched at 30 deg. angle of attack and then forced to oscillate in roll around the symmetry axis at a constant amplitude of 4 deg. and reduced frequency of pi and 2(pi). The freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. This multidisciplinary problem is solved using three sets of equations on a dynamic multi-block grid structure. The first set is the unsteady, full Navier-Stokes equations, the second set is the aeroelastic equations for coupled bending and torsion vibrations of the tails, and the third set is the grid-displacement equations. The configuration is investigated for inboard position of the twin tails which corresponds to a separation distance between the twin tails of 33% wing span. The computed results are compared with the results of stationary configuration, which previously have been validated using experimental data. The results conclusively showed that the rolling oscillations of the configuration have led to higher loads, higher deflections, and higher excitation peaks than those of the stationary configuration. Moreover, increasing the reduced frequency has led to higher loads and excitation peaks and lower bending and torsion deflections and acceleration.

Effects of Passive Porosity on Interacting Vortex Flows At Supersonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2000-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity on vortex flow interaction about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used and included pressure-sensitive paint (PSP), schlieren, and laser vapor screen (LVS) These techniques were combined with force and moment and conventional electronically-scanned pressure (ESP) measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model.

Wind tunnel investigation of the interaction and breakdown characteristics of slender wing vortices at subsonic, transonic, and supersonic speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

1991-01-01

The vortex dominated aerodynamic characteristics of a generic 65 degree cropped delta wing model were studied in a wind tunnel at subsonic through supersonic speeds. The lee-side flow fields over the wing-alone configuration and the wing with leading edge extension (LEX) added were observed at M (infinity) equals 0.40 to 1.60 using a laser vapor screen technique. These results were correlated with surface streamline patterns, upper surface static pressure distributions, and six-component forces and moments. The wing-alone exhibited vortex breakdown and asymmetry of the breakdown location at the subsonic and transonic speeds. An earlier onset of vortex breakdown over the wing occurred at transonic speeds due to the interaction of the leading edge vortex with the normal shock wave. The development of a shock wave between the vortex and wing surface caused an early separation of the secondary boundary layer. With the LEX installed, wing vortex breakdown asymmetry did not occur up to the maximum angle of attack in the present test of 24 degrees. The favorable interaction of the LEX vortex with the wing flow field reduced the effects of shock waves on the wing primary and secondary vortical flows. The direct interaction of the wing and LEX vortex cores diminished with increasing Mach number. The maximum attainable vortex-induced pressure signatures were constrained by the vacuum pressure limit at the transonic and supersonic speeds.

Aerodynamic control of NASP-type vehicles through vortex manipulation. Volume 3: Wing rock experiments

NASA Technical Reports Server (NTRS)

Suarez, Carlos J.; Smith, Brooke C.; Kramer, Brian R.; Ng, T. Terry; Ong, Lih-Yenn; Malcolm, Gerald N.

1993-01-01

Free-to-roll tests were conducted in water and wind tunnels in an effort to investigate the mechanisms of wing rock on a NASP-type vehicle. The configuration tested consisted of a highly-slender forebody and a 78 deg swept delta wing. In the water tunnel test, extensive flow visualization was performed and roll angle histories were obtained. In the wind tunnel test, the roll angle, forces and moments, and limited forebody and wing surface pressures were measured during the wing rock motion. A limit cycle oscillation was observed for angles of attack between 22 deg and 30 deg. In general, the experiments confirmed that the main flow phenomena responsible for the wing-body-tail wing rock are the interactions between the forebody and the wing vortices. The variation of roll acceleration (determined from the second derivative of the roll angle time history) with roll angle clearly slowed the energy balance necessary to sustain the limit cycle oscillation. Different means of suppressing wing rock by controlling the forebody vortices using small blowing jets were also explored. Steady blowing was found to be capable of suppressing wing rock, but significant vortex asymmetrices are created, causing the model to stop at a non-zero roll angle. On the other hand, alternating pulsed blowing on the left and right sides of the fore body was demonstrated to be a potentially effective means of suppressing wing rock and eliminating large asymmetric moments at high angles of attack.

Experimental Investigation of the Influence of a Reverse Delta Type Add-on Device on the Flap-tip Vortex of a Wing

NASA Astrophysics Data System (ADS)

Altaf, A.; Thong, T. B.; Omar, A. A.; Asrar, W.

2017-03-01

Particle Image Velocimetry was used in a low speed wind tunnel to investigate the effect of interactions of vortices produced by an outboard flap-tip of a half wing (NACA 23012 in landing configuration) and a slender reverse delta type add-on device, placed in the proximity of the outboard flap-tip, on the upper surface of the half wing. This work investigates the characteristics of the vortex interactions generated downstream in planes perpendicular to the free stream direction at a chord-based Reynolds number of Rec=2.74×105 . It was found that the add-on device significantly reduces the tangential velocity magnitude and enlarges the vortex core of the resultant vortex by up to 36.1% and 36.8%, respectively.

Effects of Various Fillet Shapes on a 76/40 Double Delta Wing from Mach 0.18 to 0.7

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Bell, James H.; Gonzalez, Hugo A.; McLachlan, Blair G.

2003-01-01

The effects of linear, diamond, and parabolic fillets on a double delta wing were investigated in the NASA Langley 7 x 10 ft High Speed Tunnel from Mach 0.18 to 0.7 and angles of attack from 4 deg. to 42 deg. Force and moment, pneumatic pressures, pressure sensitive paint, and vapor screen flow visualization measurements were used to characterize the flow field and to determine longitudinal forces and moments. The fillets increased lift coefficient and reduced induced drag without significantly affecting pitching moment. Pressure sensitive paint showed the increase in lift is caused by an increase in suction and broadening of the vortex suction footprint. Vapor screen results showed the mixing and coalescing of the strake fillet and wing vortices causes the footprint to broaden.

Force and moment measurements on a 74 deg delta wing with an apex flap

NASA Technical Reports Server (NTRS)

Buter, T. A.; Rao, D. M.

1984-01-01

Results are presented of a subsonic experimental investigation of an apex flap concept on a 74 deg swept delta wing with trailing-edge flaps. The apex flap comprised approximately 6 percent of the wing area forward of a transverse hinge, allowing for upward and downward deflection angles from +40 deg to -20 deg. Upward deflection forces leading-edge vortex formation on the apex flap, resulting in an increased lift component on the apex area. The associated nose-up moment balances the nose-down moment due to trailing-edge flaps, resulting in sizeable increase in the trimmed lift coefficient particularly at low angles of attack. Nose-down apex deflection may be used to augment the pitch control for rapid recovery from high-alpha maneuvers. This report presents the balance data without analysis.

Subsonic investigations of vortex interaction control for enhanced high-alpha aerodynamics of a chine forebody/Delta wing configuration

NASA Technical Reports Server (NTRS)

Rao, Dhanvada M.; Bhat, M. K.

1992-01-01

A proposed concept to alleviate high alpha asymmetry and lateral/directional instability by decoupling of forebody and wing vortices was studied on a generic chine forebody/ 60 deg. delta configuration in the NASA Langley 7 by 10 foot High Speed Tunnel. The decoupling technique involved inboard leading edge flaps of varying span and deflection angle. Six component force/moment characteristics, surface pressure distributions and vapor-screen flow visualizations were acquired, on the basic wing-body configuration and with both single and twin vertical tails at M sub infinity = 0.1 and 0.4, and in the range alpha = 0 to 50 deg and beta = -10 to +10 degs. Results are presented which highlight the potential of vortex decoupling via leading edge flaps for enhanced high alpha lateral/directional characteristics.

77 FR 13187 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2012-03-06

... butt splices, wing-to-body fairing locations, and external repair and cutout reinforcement areas; and... splices, wing-to-body fairing locations, and external repair and cutout reinforcement areas; and related... and Delta Airlines requested that we revise paragraph (g) of the SNPRM (75 FR 52907, August 30, 2010...

5 CFR Appendix C to Subpart B of... - Appropriated Fund Wage and Survey Areas

Code of Federal Regulations, 2011 CFR

2011-01-01

... Delta Dolores Fremont Gunnison Hinsdale Huerfano Kiowa Kit Carson Las Animas Lincoln Mineral Montrose...: Lucas Wood Northwestern Michigan Survey Area Michigan: Delta Dickinson Marquette Area of Application... Earth Indian Reservation portion only) Beltrami Cass Clearwater Cook Crow Wing Hubbard Itasca...

5 CFR Appendix C to Subpart B of... - Appropriated Fund Wage and Survey Areas

Code of Federal Regulations, 2010 CFR

2010-01-01

... Delta Dolores Fremont Gunnison Hinsdale Huerfano Kiowa Kit Carson Las Animas Lincoln Mineral Montrose...: Lucas Wood Northwestern Michigan Survey Area Michigan: Delta Dickinson Marquette Area of Application... Earth Indian Reservation portion only) Beltrami Cass Clearwater Cook Crow Wing Hubbard Itasca...

A Discrete-Vortex Method for Studying the Wing Rock of Delta Wings

NASA Technical Reports Server (NTRS)

Gainer, Thomas G.

2002-01-01

A discrete-vortex method is developed to investigate the wing rock problem associated with highly swept wings. The method uses two logarithmic vortices placed above the wing to represent the vortex flow field and uses boundary conditions based on conical flow, vortex rate of change of momentum, and other considerations to position the vortices and determine their strengths. A relationship based on the time analogy and conical-flow assumptions is used to determine the hysteretic positions of the vortices during roll oscillations. Static and dynamic vortex positions and wing rock amplitudes and frequencies calculated by using the method are generally in good agreement with available experimental data. The results verify that wing rock is caused by hysteretic deflections of the vortices and indicate that the stabilizing moments that limit wing rock amplitudes are the result of the one primary vortex moving outboard of the wing where it has little influence on the wing.

Prediction and control of vortex-dominated and vortex-wake flows

NASA Technical Reports Server (NTRS)

Kandil, Osama

1993-01-01

This progress report documents the accomplishments achieved in the period from December 1, 1992 until November 30, 1993. These accomplishments include publications, national and international presentations, NASA presentations, and the research group supported under this grant. Topics covered by documents incorporated into this progress report include: active control of asymmetric conical flow using spinning and rotary oscillation; supersonic vortex breakdown over a delta wing in transonic flow; shock-vortex interaction over a 65-degree delta wing in transonic flow; three dimensional supersonic vortex breakdown; numerical simulation and physical aspects of supersonic vortex breakdown; and prediction of asymmetric vortical flows around slender bodies using Navier-Stokes equations.

Subsonic balance and pressure investigation of a 60-deg delta wing with leading-edge devices (data report)

NASA Technical Reports Server (NTRS)

Rao, D. M.; Tingas, S. A.

1981-01-01

The drag reduction potential of leading edge devices on a 60 degree delta wing at high lift was examined. Geometric variations of fences, chordwise slots, pylon type vortex generators, leading edge vortex flaps, and sharp leading edge extensions were tested individually and in specific combinations to improve high-alpha drag performance with a minimum of low-alpha drag penalty. The force, moment, and surface static pressure data for angles of attack up to 23 degrees, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter are documented.

Computational Test Cases for a Clipped Delta Wing with Pitching and Trailing-Edge Control Surface Oscillations

NASA Technical Reports Server (NTRS)

Bennett, Robert M.; Walker, Charlotte E.

1999-01-01

Computational test cases have been selected from the data set for a clipped delta wing with a six-percent-thick circular-arc airfoil section that was tested in the NASA Langley Transonic Dynamics Tunnel. The test cases include parametric variation of static angle of attack, pitching oscillation frequency, trailing-edge control surface oscillation frequency, and Mach numbers from subsonic to low supersonic values. Tables and plots of the measured pressures are presented for each case. This report provides an early release of test cases that have been proposed for a document that supplements the cases presented in AGARD Report 702.

Correlation of Predicted and Flight Derived Stability and Control Derivatives with Particular Application to Tailless Delta Wing Configurations

NASA Technical Reports Server (NTRS)

Weil, J.

1981-01-01

Flight derived longitudinal and lateral-directional stability and control derivatives were compared to wind-tunnel derived values. As a result of these comparisons, boundaries representing the uncertainties that could be expected from wind-tunnel predictions were established. These boundaries provide a useful guide for control system sensitivity studies prior to flight. The primary application for this data was the space shuttle, and as a result the configurations included in the study were those most applicable to the space shuttle. The configurations included conventional delta wing aircraft as well as the X-15 and lifting body vehicles.

75 FR 2541 - Environmental Impact Statements and Regulations; Availability of EPA Comments

Federal Register 2010, 2011, 2012, 2013, 2014

2010-01-15

..., 158th Fighter Wing Vermont Air National Guard Project, Proposed Realignment of National Guard Avenue and... the preparing agency. EIS No. 20090401, ERP No. F-IBR-K39118-CA, Delta-Mendota Canal/ California... Delta-Mendota Water Authority Project, Central Valley Project, Alameda and San Joaquin Counties, CA...

Investigation at transonic speeds of the lateral-control and hinge-moment characteristics of a flap-type spoiler aileron on a 60 degree delta wing

NASA Technical Reports Server (NTRS)

Wiley, Harleth G; Taylor, Robert T

1954-01-01

This paper present results of an investigation of the lateral-control and hinge-moment characteristics of a 0.67 semispan flap-type spoiler aileron on a semispan thin 60 degree delta wing at transonic speeds by the reflection-plane technique. The spoiler-aileron had a constant chord of 10.29 percent mean aerodynamic chord and was hinged at the 81.9-percent-wing-root-chord station. Tests were made with the spoiler aileron slot open, partially closed, and closed. Incremental rolling-moment coefficients were obtained through a Mach number range of 0.62 to 1.08. Results indicated reasonably linear variations of rolling-moment and hinge-moment coefficients with spoiler projection except at spoiler projections of less than -2 percent mean aerodynamic chord and angles of attack greater than 12 degrees with results generally independent of slot geometry.

A Computational and Experimental Investigation of a Delta Wing with Vertical Tails

NASA Technical Reports Server (NTRS)

Krist. Sherrie L.; Washburn, Anthony E.; Visser, Kenneth D.

2004-01-01

The flow over an aspect ratio 1 delta wing with twin vertical tails is studied in a combined computational and experimental investigation. This research is conducted in an effort to understand the vortex and fin interaction process. The computational algorithm used solves both the thin-layer Navier-Stokes and the inviscid Euler equations and utilizes a chimera grid-overlapping technique. The results are compared with data obtained from a detailed experimental investigation. The laminar case presented is for an angle of attack of 20 and a Reynolds number of 500; 000. Good agreement is observed for the physics of the flow field, as evidenced by comparisons of computational pressure contours with experimental flow-visualization images, as well as by comparisons of vortex-core trajectories. While comparisons of the vorticity magnitudes indicate that the computations underpredict the magnitude in the wing primary-vortex-core region, grid embedding improves the computational prediction.

A computational and experimental investigation of a delta wing with vertical tails

NASA Technical Reports Server (NTRS)

Krist, Sherrie L.; Washburn, Anthony E.; Visser, Kenneth D.

1993-01-01

The flow over an aspect ratio 1 delta wing with twin vertical tails is studied in a combined computational and experimental investigation. This research is conducted in an effort to understand the vortex and fin interaction process. The computational algorithm used solves both the thin-layer Navier-Stokes and the inviscid Euler equations and utilizes a chimera grid-overlapping technique. The results are compared with data obtained from a detailed experimental investigation. The laminar case presented is for an angle of attack of 20 deg and a Reynolds number of 500,000. Good agreement is observed for the physics of the flow field, as evidenced by comparisons of computational pressure contours with experimental flow-visualization images, as well as by comparisons of vortex-core trajectories. While comparisons of the vorticity magnitudes indicate that the computations underpredict the magnitude in the wing primary-vortex-core region, grid embedding improves the computational prediction.

Aerodynamic stability and control characteristics of TBC shuttle booster AR-11981-3

NASA Technical Reports Server (NTRS)

Phelps, E. R.; Watts, L. L.; Ainsworth, R. W.

1972-01-01

A scale model of the Boeing Company space shuttle booster configuration 3 was tested in the MSFC 14-inch trisonic wind tunnel. This test was proposed to fill-in the original test run schedule as well as to investigate the aerodynamic stability and control characteristics of the booster with three wing configurations not previously tested. The configurations tested included: (1) a cylindrical booster body with an axisymmetric nose, (2) clipped delta canards that had variable incidence from 0 deg to -60 deg, (3) different aft body mounted wing configurations, (4) two vertical fin configurations, and (5) a Grumman G-3 orbiter configuration. Tests were conducted over a Mach range from 0.6 to 5.0.

Space shuttle: Static stability and control investigation of the NR-GD/C delta wing booster (B-15B-1) and a Reusable Nuclear Stage (RNS) M = 0.6 - 4.96

NASA Technical Reports Server (NTRS)

Allen, E. C.

1971-01-01

Experimental aerodynamic investigations were made in the NASA/MSFC 14 x 14 inch trisonic wind tunnel on a NR-GD/C 0.0031 scale model delta wing booster (B-15B-1) and reusable nuclear stage (RNS). Three basic configurations were tested: (1) the B-15B-1 booster alone, (2) the RNS alone, and (3) the booster with the RNS mounted piggyback. Six component force and moment data were recorded for each of these configurations over an angle of attack range from -16 deg to 4 deg at zero degrees sideslip, and over an angle of sideslip range from -10 deg to 10 deg at zero and -6 degrees angle of attack. The configurations were tested over a Mach number range of 0.6 to 5.0 with a nominal Reynolds number of 6.5 million per foot except for Mach 1.2 and 3.0 where the Reynolds number were 12.4 million and 4.4 million per foot respectively.

Pitot-pressure distributions of the flow field of a delta-wing orbiter

NASA Technical Reports Server (NTRS)

Cleary, J. W.

1972-01-01

Pitot pressure distributions of the flow field of a 0.0075-scale model of a typical delta wing shuttle orbiter are presented. Results are given for the windward and leeward sides on centerline in the angle-of-attack plane from wind tunnel tests conducted in air. Distributions are shown for three axial stations X/L = .35, .60, and .98 and for angles of attack from 0 to 60 deg. The tests were made at a Mach number of 7.4 and for Reynolds numbers based on body length from 1,500,000 to 9,000,000. The windward distributions at the two survey stations forward of the body boat tail demonstrate the compressive aspects of the flow from the shock wave to the body. Conversely, the distributions at the aft station display an expansion of the flow that is attributed to body boat tail. On the lee side, results are given at low angles of attack that illustrate the complicating aspects of the canopy on the flow field, while results are given to show the effects of flow separation at high angles of attack.

Wing Wake Vortices and Temporal Vortex Pair Instabilities

NASA Astrophysics Data System (ADS)

Williamson, C. H. K.; Leweke, T.; Miller, G. D.

In this presentation we include selected results which have originated from vortex dynamics studies conducted at Cornell, in collaboration with IRPHE, Marseille. These studies concern, in particular, the spatial development of delta wing trailing vortices, and the temporal development of counter-rotating vortex pairs. There are, as might be expected, similarities in the instabilities of both of these basic flows, as shown in our laboratory-scale studies. In the case of the spatial development of vortex pairs in the wake of a delta wing, either in free flight or towed from an XY carriage system in a towing tank, we have found three distinct instability length scales as the trailing vortex pair travels downstream. The first (smallest-scale) instability is found immediately behind the delta wing, and this scales on the thickness of the two shear layers separating from the wing trailing edge. The second (short-wave) instability, at an intermediate distance downstream, scales on the primary vortex core dimensions. The third (long-wave) instability far downstream represents the classical "Crow" instability (Crow, 1970), scaling on the distance between the two primary vortices. By imposing disturbances on the delta wing incident velocity, we find that the long-wave instability is receptive to a range of wavelengths. Our experimental measurements of instability growth rates are compared with theoretical predictions, which are based on the theory of Widnall et al. (1971), and which require, as input, DPIV measurements of axial and circumferential velocity profiles. This represents the first time that theoretical and experimental growth rates have been compared, without the imposition of ad-hoc assumptions regarding the vorticity distribution. The agreement with theory appears to be good. The ease with which a Delta wing may be flown in free flight was demonstrated at the Symposium, using a giant polystyrene triangular wing, launched from the back of the auditorium, and ably caught by Professor Sid Leibovich, in whose honour the Symposium was held. In the case of the temporal growth of vortex pairs, formed by the closing of a pair of long flaps underwater, we find two principal instabilities; namely, a longwavelength Crow instability, and a short-wavelength "elliptic" instability. Comparisons between experiment and theory for the growth rates of the long-wave instability, over a range of perturbed wavelengths, appears to be very good. The vortex pair "pinches off", or reconnects, to form vortex rings in the manner assumed to occur in contrails behind jet aircraft. We discover a symmetry-breaking phase relationship for the short wave disturbances growing in the two vortices, which we 380 C.H.K. Williamson et al. show to be consistent with a kinematic matching condition between the two disturbances. Further results demonstrate that this instability is a manifestation of an elliptic instability, which is here identified for the first time in a real open flow. We therefore refer to this flow as a "cooperative elliptic" instability. The long-term evolution of the flow involves the inception of secondary miniscule vortex pairs, which are perpendicular to the primary vortex pair.

Effects of forebody strakes and Mach number on overall aerodynamic characteristics of configuration with 55 deg cropped delta wing

NASA Technical Reports Server (NTRS)

Erickson, Gary E.; Rogers, Lawrence W.

1992-01-01

A wind tunnel data base was established for the effects of chine-like forebody strakes and Mach number on the longitudinal and lateral-directional characteristics of a generalized 55 degree cropped delta wing-fuselage-centerline vertical tail configuration. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center at free-stream Mach numbers of 0.40 to 1.10 and Reynolds numbers based on the wing mean aerodynamic chord of 1.60 x 10(exp 6) to 2.59 x 10(exp 6). The best matrix included angles of attack from 0 degree to a maximum of 28 degree, angles of sidesip of 0, +5, and -5 degrees, and wing leading-edge flat deflection angles of 0 and 30 degrees. Key flow phenomena at subsonic and transonic conditions were identified by measuring off-body flow visualization with a laser screen technique. These phenomena included coexisting and interacting vortex flows and shock waves, vortex breakdown, vortex flow interactions with the vertical tail, and vortices induced by flow separation from the hinge line of the deflected wing flap. The flow mechanisms were correlated with the longitudinal and lateral-directional aerodynamic data trends.

Research in computational fluid dynamics

NASA Technical Reports Server (NTRS)

Murman, Earll M.

1987-01-01

The numerical integration of quasi-one-dimensional unsteady flow problems which involve finite rate chemistry are discussed, and are expressed in terms of conservative form Euler and species conservation equations. Hypersonic viscous calculations for delta wing geometries is also examined. The conical Navier-Stokes equations model was selected in order to investigate the effects of viscous-inviscid interations. The more complete three-dimensional model is beyond the available computing resources. The flux vector splitting method with van Leer's MUSCL differencing is being used. Preliminary results were computed for several conditions.

Some effects of wing and body geometry on the aerodynamic characteristics of configurations designed for high supersonic Mach numbers

NASA Technical Reports Server (NTRS)

Spearman, M. L.; Tice, David C.; Braswell, Dorothy O.

1992-01-01

Experimental and theoretical results are presented for a family of aerodynamic configurations for flight Mach numbers as high as Mach 8. All of these generic configurations involved 70-deg sweep delta planform wings of three different areas and three fuselage shapes with circular-to-elliptical cross sections. It is noted that fuselage ellipticity enhances lift-curve slope and maximum L/D, while decreasing static longitudinal stability (especially with smaller wing areas).

Prediction de l’Eclatement Tourbillonnaire sur les Ailes Delta d’Avions Militaires (CFD Prediction of Vortex Breakdown on Delta Wings for Military Aircraft)

DTIC Science & Technology

2003-03-01

moins développé que le tourbillon expérimental. En moyenne, par effet de compensation entre la région de bord d’attaque et la région proche de ...que les effets d’anisotropie de la turbulence. Un premier élément de réponse peut être obtenu en utilisant un modèle de turbulence un plus élaboré, tel...que l’EARSM (Explicit Algebraïc Reynolds Stress Model) qui prend en compte les effets de rotation et d’anisotropie [Ref 6]. Nous verrons au paragraphe

78 FR 61621 - Endangered and Threatened Wildlife and Plants; Proposed Threatened Status for the Western...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-10-03

... length, depth of upper mandible, and wing length) of almost 700 specimens of adult yellow-billed cuckoos... wing length, bill length, and mandible depth; and (3) no statistically significant differences for... west side of the Cascades at the Sandy River Delta near its confluence with the Columbia River in July...

Overview of Unsteady Transonic Wind Tunnel Test on a Semispan Straked Delta Wing Oscillating in Pitch

DTIC Science & Technology

1994-08-01

of research between the Lockheed Fort Worth Company (LFWC), Fort Worth, Texas, USA (Formerly the Fort Worth Division of General Dynamics until 28...Separation," NASA CR 4090, August 1987. 13. Cunningham, A.M., Jr. and Bushlow: "Steady and Unsteady Force Testing of Fighter Aircraft Models in a Water Tunnel," AIAA Papaer No. AIAA-90-2815, August 1990. 47

A Preliminary Analysis of the Flying Qualities of the Consolidated Vultee MX-813 Delta-Wing Airplane Configuration at Transonic and Low Supersonic Speeds as Determined from Flights of Rocket-Powered Models

NASA Technical Reports Server (NTRS)

Mitcham, Grady L.

1949-01-01

A preliminary analysis of the flying qualities of the Consolidated Vultee MX-813 delta-wing airplane configuration has been made based on the results obtained from the first two 1/8 scale models flown at the NACA Pilotless Aircraft Research Station, Wallop's Island, VA. The Mach number range covered in the tests was from 0.9 to 1.2. The analysis indicates adequate elevator control for trim in level flight over the speed range investigated. Through the transonic range there is a mild trim change with a slight tucking-under tendency. The elevator control effectiveness in the supersonic range is reduced to about one-half the subsonic value although sufficient control for maneuvering is available as indicated by the fact that 10 deg elevator deflection produced 5g acceleration at Mach number of 1.2 at 40,000 feet.The elevator control forces are high and indicate the power required of the boost system. The damping. of the short-period oscillation is adequate at sea-level but is reduced at 40,000 feet. The directional stability appears adequate for the speed range and angles of attack covered.

Development and demonstration of a flutter-suppression system using active controls. [wind tunnel tests

NASA Technical Reports Server (NTRS)

Sandford, M. C.; Abel, I.; Gray, D. L.

1975-01-01

The application of active control technology to suppress flutter was demonstrated successfully in the transonic dynamics tunnel with a delta-wing model. The model was a simplified version of a proposed supersonic transport wing design. An active flutter suppression method based on an aerodynamic energy criterion was verified by using three different control laws. The first two control laws utilized both leading-edge and trailing-edge active control surfaces, whereas the third control law required only a single trailing-edge active control surface. At a Mach number of 0.9 the experimental results demonstrated increases in the flutter dynamic pressure from 12.5 percent to 30 percent with active controls. Analytical methods were developed to predict both open-loop and closed-loop stability, and the results agreed reasonably well with the experimental results.

Vortex boundary-layer interactions

NASA Technical Reports Server (NTRS)

Bradshaw, P.

1986-01-01

Parametric studies to identify a vortex generator were completed. Data acquisition in the first chosen configuration, in which a longitudinal vortex pair generated by an isolated delta wing starts to merge with a turbulent boundary layer on a flat plate fairly close to the leading edge is nearly completed. Work on a delta-wing/flat-plate combination, consisting of a flow visualization and hot wire measurements taken with a computer controlled traverse gear and data logging system were completed. Data taking and analysis have continued, and sample results for another cross stream plane are presented. Available data include all mean velocity components, second order mean products of turbulent fluctuations, and third order mean products. Implementation of a faster data logging system was accomplished.

Numerical Simulation of Forced and Free-to-Roll Delta-Wing Motions

NASA Technical Reports Server (NTRS)

Chaderjian, Neal M.; Schiff, Lewis B.

1996-01-01

The three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate nonsteady vortical flow about a 65-deg sweep delta wing at 30-deg angle of attack. Two large-amplitude, high-rate, forced-roll motions, and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are in good agreement with the forces, moments, and roll-angle time histories. Vortex breakdown is present in each case. Significant time lags in the vortex breakdown motions relative to the body motions strongly influence the dynamic forces and moments.

Wing rock suppression using forebody vortex control

NASA Technical Reports Server (NTRS)

Ng, T. T.; Ong, L. Y.; Suarez, C. J.; Malcolm, G. N.

1991-01-01

Static and free-to-roll tests were conducted in a water tunnel with a configuration that consisted of a highly-slender forebody and 78-deg sweep delta wings. Flow visualization was performed and the roll angle histories were obtained. The fluid mechanisms governing the wing rock of this configuration were identified. Different means of suppressing wing rock by controlling the forebody vortices using small blowing jets were also explored. Steady blowing was found to be capable of suppressing wing rock, but significant vortex asymmetries had to be induced at the same time. On the other hand, alternating pulsed blowing on the left and right sides of the forebody was demonstrated to be potentially an effective means of suppressing wing rock and eliminating large asymmetric moments at high angles of attack.

Types of flow on the lee side of delta wings

NASA Astrophysics Data System (ADS)

Narayan, K. Yegna; Seshadri, S. N.

1997-03-01

Delta wings have found wide application in a variety of aerospace vehicles including high performance combat aircraft, supersonic civil aircraft, (proposed) hypersonic aircraft and the space shuttle orbiter. A considerable amount of research work has been carried out over the past three decades and an extensive body of literature is available. The present review focuses attention on the nine possible types of flow that can occur on the lee side of delta wings in a Mach number range which extends from subsonic to hypersonic. The dependence of the flow types on geometrical and freestream parameters has been discussed in detail. The extensive experimental data available has made it possible to obtain a broad physical understanding of the mechanisms underlying the different flow types. However much more work needs to be done to determine the effects of Reynolds number, particularly when either the state of the boundary layer is transitional or when the type of flow is changing from leading edge attached to separated. Computational methods have made spectacular advances in recent years. In particular, solutions of Reynolds averaged Navier-Stokes equations at fairly high Reynolds number have become possible and these computations have captured eight of the nine experimentally observed flow types, including those involving cross flow shock waves and shock-induced separation.

Flow solution on a dual-block grid around an airplane

NASA Technical Reports Server (NTRS)

Eriksson, Lars-Erik

1987-01-01

The compressible flow around a complex fighter-aircraft configuration (fuselage, cranked delta wing, canard, and inlet) is simulated numerically using a novel grid scheme and a finite-volume Euler solver. The patched dual-block grid is generated by an algebraic procedure based on transfinite interpolation, and the explicit Runge-Kutta time-stepping Euler solver is implemented with a high degree of vectorization on a Cyber 205 processor. Results are presented in extensive graphs and diagrams and characterized in detail. The concentration of grid points near the wing apex in the present scheme is shown to facilitate capture of the vortex generated by the leading edge at high angles of attack and modeling of its interaction with the canard wake.

Vortex flow hysteresis

NASA Technical Reports Server (NTRS)

Cunningham, A. M., Jr.

1986-01-01

An experimental study was conducted to quantify the hysteresis associated with various vortex flow transition points and to determine the effect of planform geometry. The transition points observed consisted of the appearance (or disappearance) of trailing edge vortex burst and the transition to (or from) flat plate or totally separated flows. Flow visualization with smoke injected into the vortices was used to identify the transitions on a series of semi-span models tested in a low speed tunnel. The planforms tested included simple deltas (55 deg to 80 deg sweep), cranked wings with varying tip panel sweep and dihedral, and a straked wing. High speed movies at 1000 frames per second were made of the vortex flow visualization in order to better understand the dynamics of vortex flow, burst and transition.

18 CFR Appendix A to Part 11 - Fee Schedule for FY 2009

Code of Federal Regulations, 2010 CFR

2010-04-01

... Delta $46.21 Colorado Denver * $23.10 Colorado Dolores $23.10 Colorado Douglas $69.31 Colorado Eagle $34... $46.21 Michigan Crawford $69.31 Michigan Delta $34.66 Michigan Dickinson $34.66 Michigan Eaton $69.31....10 Minnesota Cook $34.66 Minnesota Cottonwood $34.66 Minnesota Crow Wing $23.10 Minnesota Dakota $69...

18 CFR Appendix A to Part 11 - Fee Schedule for FY 2010

Code of Federal Regulations, 2011 CFR

2011-04-01

... Custer 47.08 Colorado Delta 62.78 Colorado Denver * 31.39 Colorado Dolores 31.39 Colorado Douglas 94.17... Michigan Delta 47.08 Michigan Dickinson 47.08 Michigan Eaton 94.17 Michigan Emmet 94.17 Michigan Genesee... Minnesota Cook 47.08 Minnesota Cottonwood 47.08 Minnesota Crow Wing 31.39 Minnesota Dakota 94.17 Minnesota...

75 FR 1447 - Alaska Railroad Corporation-Construction and Operation Exemption-Rail Line Between North Pole and...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-01-11

... Railroad Corporation--Construction and Operation Exemption--Rail Line Between North Pole and Delta Junction... of North Pole (located just south of Fairbanks) to the southern side of the community of Delta... Railroad Administration, U.S. Air Force 354th Fighter Wing Command from Eielson Air Force Base, U.S. Army...

78 FR 14644 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-03-07

... within the ECS bay, which in combination with flammable fuel vapors, could result in a center wing fuel... ECS bay, but allows installation of clamp P/N TA0930034-11 at the same clamp position. Delta Air Lines (Delta) requested that we ensure that paragraph (h) of the NPRM, only applies to those airplanes subject...

Definition of the unsteady vortex flow over a wing/body configuration

NASA Technical Reports Server (NTRS)

Liou, S. G.; Debry, B.; Lenakos, J.; Caplin, J.; Komerath, N. M.

1991-01-01

A problem of current interest in computational aerodynamics is the prediction of unsteady vortex flows over aircraft at high angles of attack. A six-month experimental effort was conducted at the John H. Harper Wind Tunnel to acquire qualitative and quantitative information on the unsteady vortex flow over a generic wing-body configuration at high angles of attack. A double-delta flat-plate wing with beveled edges was combined with a slender sharp-nosed body-of-revolution fuselage to form the generic configuration. This configuration produces a strong attached leading edge vortex on the wing, as well as sharply-peaked flow velocity spectra above the wing. While it thus produces flows with several well-defined features of current interest, the model was designed for efficiency of representation in computational codes. A moderate number of surface pressure ports and two unsteady pressure sensors were used to study the pressure distribution over the wing and body surface at high angles of attack; the unsteady pressure sensing did not succeed because of inadequate signal-to-noise ratio. A pulsed copper vapor laser sheet was used to visualize the vortex flow over the model, and vortex trajectories, burst locations, mutual induction of vortex systems from the forebody, strake, and wing, were quantified. Laser Doppler velocimetry was used to quantify all 3 components of the time-average velocity in 3 data planes perpendicular to the freestream direction. Statistics of the instantaneous velocity were used to study intermittency and fluctuation intensity. Hot-film anemometry was used to study the fluctuation energy content in the velocity field, and the spectra of these fluctuations. In addition, a successful attempt was made to measure velocity spectra, component by component, using laser velocimetry, and these were compared with spectra measured by hot-film anemometry at several locations.

76 FR 26655 - Public Meetings on the Proposed Rule for Mandatory Inspection of Catfish and Catfish Products

Federal Register 2010, 2011, 2012, 2013, 2014

2011-05-09

... photo ID to enter the building. The Jefferson Auditorium is located at Wing 6 in the South Building. Attendees should enter the building via Wing 5 or 7 on 14th Street and Independence Avenue, SW. The second... at the Delta Research and Extension Center of the Mississippi State University. The Charles Capp...

77 FR 12448 - Airworthiness Directives; Pratt & Whitney Division Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-03-01

... Delta Airlines, Inc. requested that we specify replacing affected FMUs with FMU P/N 53U044, or later FAA... actual cost of the AD; not other costs. We did not change the AD. Request To Replace FMUs On-Wing United Parcel Service Co. requested that we include in the AD the option to perform on-wing replacements of...

14 CFR Special Federal Aviation... - Mitsubishi MU-2B Series Special Training, Experience, and Operating Requirements

Code of Federal Regulations, 2010 CFR

2010-01-01

.... Fuselage 2. Wing 3. Empennage 4. Doors 5. Windshield and Windows C. Airplane Systems 1. Electrical Power 2... System 1. Navigation Lights 2. Anti-Collision Lights 3. Wing Inspection Lights 4. Taxi Lights 5. Landing... Operations 3. Synchronization 4. De-Ice D. Ground Checks 1. Overspeed Governor 2. SRL and Delta P/P 3. NTS...

14 CFR Special Federal Aviation... - Mitsubishi MU-2B Series Special Training, Experience, and Operating Requirements

Code of Federal Regulations, 2011 CFR

2011-01-01

.... Fuselage 2. Wing 3. Empennage 4. Doors 5. Windshield and Windows C. Airplane Systems 1. Electrical Power 2... System 1. Navigation Lights 2. Anti-Collision Lights 3. Wing Inspection Lights 4. Taxi Lights 5. Landing... Operations 3. Synchronization 4. De-Ice D. Ground Checks 1. Overspeed Governor 2. SRL and Delta P/P 3. NTS...

Lee-side flow over delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Miller, D. S.; Wood, R. M.

1985-01-01

An experimental investigation of the lee-side flow on sharp leading-edge delta wings at supersonic speeds has been conducted. Pressure data were obtained at Mach numbers from 1.5 to 2.8, and three types of flow-visualization data (oil-flow, tuft, and vapor-screen) were obtained at Mach numbers from 1.7 to 2.8 for wing leading-edge sweep angles from 52.5 deg to 75 deg. From the flow-visualization data, the lee-side flows were classified into seven distinct types and a chart was developed that defines the flow mechanism as a function of the conditions normal to the wing leading edge, specifically, angle of attack and Mach number. Pressure data obtained experimentally and by a semiempirical prediction method were employed to investigate the effects of angle of attack, leading-edge sweep, and Mach number on vortex strength and vortex position. In general, the predicted and measured values of vortex-induced normal force and vortex position obtained from experimental data have the same trends with angle of attack, Mach number, and leading-edge sweep; however, the vortex-induced normal force is underpredicted by 15 to 30 percent, and the vortex spanwise location is overpredicted by approximately 15 percent.

Morphometric Variation of the Aedes albifasciatus (Diptera: Culicidae) Wings in Three Populations From Different Ecoregions of Argentina.

PubMed

Garzón, Maximiliano J; Schweigmann, Nicolás

2018-06-23

Shape variability among individuals is important to understand some ecological relationships, since it provides the nexus between the genotype and the environment. Geometric morphometrics based on generalized procrustes analysis was applied on 17 landmarks of the wings of Aedes albifasciatus (Macquart 1838) (Diptera: Culicidae) females collected from three ecoregions of Argentina (Delta and islands of the Paraná River, Pampa, and Patagonian steppe). This methodology was used to discriminate the shapes of individuals belonging to different regions. The population of the Patagonian steppe, which was the most geographically distant, showed the most dissimilar shape. Different local variations in wing shape could have been selected according to the environmental characteristics and maintained by geographic isolation. The individuals of the two ecoregions closest to each other (Delta and islands of the Paraná River and Pampa) showed differences in shape that can be explained by a lower gene flow due to the effect of geographic isolation (by the Paraná River) and the limited dispersive capacity of Ae. albifasciatus. The results allow concluding that both environmental diversity and geographic barriers could contribute to local variations in wing shape.

Breaking down the delta wing vortex: The role of vorticity in the breakdown process. Ph.D. Thesis Final Report

NASA Technical Reports Server (NTRS)

Nelson, Robert C.; Visser, Kenneth D.

1990-01-01

Experimental x-wire measurements of the flowfield above a 70 and 75 deg flat plate delta wing were performed at a Reynolds number of 250,000. Grids were taken normal to the wing at various chordwise locations for angles of attack of 20 and 30 deg. Axial and azimuthal vorticity distributions were derived from the velocity fields. The dependence of circulation on distance from the vortex core and on chordwise location was also examined. The effects of nondimensionalization in comparison with other experimental data is made. The results indicate that the circulation distribution scales with the local semispan and grows in a nearly linear fashion in the chordwise direction. The spanwise distribution of axial vorticity is severely altered through the breakdown. The axial vorticity components with a negative sense, such as that found in the secondary vortex, seem to remain unaffected by changes in wind sweep or angle of attack, in direct contrast to the positive components. In addition, the inclusion of the local wing geometry into a previously derived correlation parameter allows the circulation of growing leading edge vortex flows to be reduced into a single curve.

Surface pressure distributions on a delta wing undergoing large amplitude pitching oscillations. M.S. Thesis

NASA Technical Reports Server (NTRS)

Thompson, Scott A.

1989-01-01

Wind tunnel experiments were performed on a 70 deg sweep delta wing to determine the effect of a sinusoidal pitching motion on the pressure field on the suction side of the wing. Twelve pressure taps were placed from 35 to 90 percent of the chord, at 60 percent of the local semi-span. Pressure coefficients were measured as a function of Reynolds number and pitch rate. The pressure coefficient was seen to vary at approximately the same frequency as the pitching frequency. The relative pressure variation at each chord location was comparable for each case. The average pressure distribution through each periodic motion was near the static distribution for the average angle of attack. Upon comparing the upstroke and downstroke pressures for a specific angle of attack, the downstroke pressures were slightly larger. Vortex breakdown was seen to have the most significant effect at the 40 to 45 percent chord location, where a decrease in pressure was apparent.

F-16XL ship #1 - CAWAP boundary layer rakes and hot film on left wing

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the boundary layer hot film and the boundary layer rakes on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 - CAWAP outboard rake #7

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the #7 outboard rake on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 wing close-up showing boundary layer detection Preston tubes

NASA Technical Reports Server (NTRS)

1995-01-01

This photo shows the boundary layer Preston tubes mounted on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Sonic-box method employing local Mach number for oscillating wings with thickness

NASA Technical Reports Server (NTRS)

Ruo, S. Y.

1978-01-01

A computer program was developed to account approximately for the effects of finite wing thickness in the transonic potential flow over an oscillating wing of finite span. The program is based on the original sonic-box program for planar wing which was previously extended to include the effects of the swept trailing edge and the thickness of the wing. Account for the nonuniform flow caused by finite thickness is made by application of the local linearization concept. The thickness effect, expressed in terms of the local Mach number, is included in the basic solution to replace the coordinate transformation method used in the earlier work. Calculations were made for a delta wing and a rectangular wing performing plunge and pitch oscillations, and the results were compared with those obtained from other methods. An input quide and a complete listing of the computer code are presented.

Development of control systems for space shuttle vehicles. Volume 2: Appendixes

NASA Technical Reports Server (NTRS)

Stone, C. R.; Chase, T. W.; Kiziloz, B. M.; Ward, M. D.

1971-01-01

A launch phase random normal wind model is presented for delta wing, two-stage, space shuttle control system studies. Equations, data, and simulations for conventional launch studies are given as well as pitch and lateral equations and data for covariance analyses of the launch phase of MSFC vehicle B. Lateral equations and data for North American 130G and 134D are also included along with a high-altitude abort simulation.

The role of flow field structure in determining the aerodynamic response of a delta wing

NASA Astrophysics Data System (ADS)

Addington, Gregory Alan

Delta wings have long been known to exhibit nonlinear aerodynamic responses as a result of the presence of helical leading-edge vortices. This nonlinearity, found under both steady-state and unsteady conditions, is particularly profound in the presence of vortex burst. Modeling such aerodynamic responses with the Nonlinear Indicial Response (NIR) methodology provides a means of simulating these nonlinearities through its inclusion of motion history in addition to superposition. The NIR model also includes provisions for a finite number of discrete locations where the aerodynamic response is discontinuous with response to a state variable. These critical states also separate regions of states where the unsteady aerodynamic responses are potentially of highly-disparate characters. Although these critical states have been found in the past, their relationship with flow field bifurcation is uncertain. The purpose of this dissertation is to explore the relationship between nonlinear aerodynamic responses, critical states and flow field bifurcations from an experimental approach. This task has been accomplished by comparing a comprehensive database of skin-friction line topologies with static and unsteady aerodynamic responses. These data were collected using a 65sp° delta wing which rolled about an inclined longitudinal body axis. In this study, compelling, but not conclusive, evidence was found to suggest that a bifurcation in the skin-friction line topology was a necessary condition for the presence of a critical state. Although the presence of critical states was well predicted through careful observation and analysis of highly-resolved static loading data alone, their precise placement as a function of the independent variable was aided through the consideration of the locations of skin-friction line bifurcations. Furthermore, these static data were found to contain indications of the basic lagged or unlagged behavior of the unsteady aerodynamic response. This indication was found by comparing the relative rate of change seen in the estimated vortical- and potential-rolling-moment components. Through the review of these data in light of current theories on the mechanisms of leading-edge vortex breakdown, the formulation of a hypothesis regarding the relationship between both the static and unsteady aerodynamic response and vorticity dynamics was possible.

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

NASA Technical Reports Server (NTRS)

Brown, Clarence A , Jr

1957-01-01

A full- scale rocket-powered model of a cruciform canard missile configuration with a low- aspect - ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed- control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift - curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift- .curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta= -0.3 deg . The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic- center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number.The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Flight Investigation at Low Angles of Attack to Determine the Longitudinal Stability and Control Characteristics of a Cruciform Canard Missile Configuration with a Low-Aspect-Ratio Wing and Blunt Nose at Mach Numbers from 1.2 to 2.1

NASA Technical Reports Server (NTRS)

Brown, C. A., Jr.

1957-01-01

A full-scale rocket-powered model of a cruciform canard missile configuration with a low-aspect-ratio wing and blunt nose has been flight tested by the Langley Pilotless Aircraft Research Division. Static and dynamic longitudinal stability and control derivatives of this interdigitated canard-wing missile configuration were determined by using the pulsed-control technique at low angles of attack and for a Mach number range of 1.2 to 2.1. The lift-curve slope showed only small nonlinearities with changes in control deflection or angle of attack but indicated a difference in lift-curve slope of approximately 7 percent for the two control deflections of delta = 3.0 deg and delta = -0.3 deg. The large tail length of the missile tested was effective in producing damping in pitch throughout the Mach number range tested. The aerodynamic-center location was nearly constant with Mach number for the two control deflections but was shown to be less stable with the larger control deflection. The increment of lift produced by the controls was small and positive throughout the Mach number range tested, whereas the pitching moment produced by the controls exhibited a normal trend of reduced effectiveness with increasing Mach number. The effectiveness of the controls in producing angle of attack, lift, and pitching moment was good at all Mach numbers tested.

Eclipse program F-106 aircraft in flight, front view

NASA Technical Reports Server (NTRS)

1997-01-01

Shot of the QF-106 aircraft in flight with the landing gear deployed. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

F-16XL ship #1 outboard rake #7

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the #7 outboard rake on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

Long-wavelength Instability of Trailing Vortices Behind a Delta Wing

NASA Astrophysics Data System (ADS)

Miller, G. D.; Williamson, C. H. K.

1996-11-01

The long-wavelength instability of a vortex pair is studied in the wake of a delta wing. While many previous studies of the instability exist, almost none are accompanied by accurate measurements of the vortex core parameters upon which the theoretical predictions depend. The present measurements of wavelength and maximum growth rate from visualization images are accompanied by extensive DPIV measurements of the distributions of vorticity and axial velocity. Axial velocity was found to be wake-like, with a velocity deficit. The vorticity distribution in the cores is well modeled by an Oseen vortex, as is the downstream growth of the core. The naturally occuring wavelength was measured to be 4.5 times the inter-vortex spacing, which compares very well with the wavelength of maximum growth rate predicted by theory using measured core parameters. Also, the measured value of the growth rate and the lower stability limit correspond well with theory. The response of the wake to forcing is also examined, and reveals that the wake is receptive to forcing at wavelengths near the natural wavelength. We demonstrate control over the rate at which the wake decays by hastening the action of the instabilty with initial forcing. Supported by NDSEG Fellowship for first author.

78 FR 8111 - Taking and Importing Marine Mammals; Taking Marine Mammals Incidental to Space Vehicle and Test...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-02-05

... hereby given that a letter of authorization (LOA) has been issued to the 30th Space Wing, U.S. Air Force...). Currently, six space launch vehicle programs use VAFB to launch satellites into polar orbit: Delta II; Taurus; Atlas V; Delta IV; Falcon; and Minotaur. Also a variety of small missiles, several types of...

77 FR 6086 - Taking and Importing Marine Mammals; Taking Marine Mammals Incidental to Space Vehicle and Test...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-02-07

... hereby given that a letter of authorization (LOA) has been issued to the 30th Space Wing, U.S. Air Force...). Currently, six space launch vehicle programs use VAFB to launch satellites into polar orbit: Delta II; Taurus; Atlas V; Delta IV; Falcon; and Minotaur. Also a variety of small missiles, several types of...

Wind-tunnel studies of the effects of simulated damage on the aerodynamic characteristics of airplanes and missiles

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1979-01-01

In order to assess the effects on static aerodynamic characteristics of battle damage to an aircraft or missile, wind tunnel studies were performed on models from which all or parts of the wing or horizontal or vertical tail had been removed. The effects of damage on the lift, longitudinal stability, lateral stability and directional stability of a swept-wing fighter are presented, along with the effects of wing removal on the control requirements of a delta-wing fighter. Results indicate that the loss of a major part of the vertical tail will probably result in the loss of the aircraft at any speed, while the loss of major parts of the horizontal tail generally results in catastrophic instability at subsonic speeds but, at low supersonic speeds, may allow the aircraft to return to friendly territory before pilot ejection. Major damage to the wing may be sustained without the loss of aircraft or pilot. The loss of some of the aerodynamic surfaces of cruise or surface-to-air missiles may result in catastrophic instability or may permit a ballistic trajectory to be maintained, depending upon the location of the lost surface with respect to the center of gravity of the missile.

Multistage aerospace craft. [perspective drawings of conceptual design

NASA Technical Reports Server (NTRS)

Kelly, D. L. (Inventor)

1973-01-01

A conceptual design of a multi-stage aerospace craft is presented. Two perspective views of the vehicle are developed to show the two component configuration with delta wing, four vertical tail surfaces, tricycle landing gear, and two rocket exhaust nozzles at the rear of the fuselage. Engines for propulsion in the atmosphere are mounted on the fuselage in front of the wing root attachment.

Wind Tunnel Investigation of Passive Porosity Applied to the Leading-Edge Extension and Leading-Edge Flaps on a Slender Wing at Subsonic Speed

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2017-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.

Determination of mean camber surfaces for wings having uniform chordwise loading and arbitrary spanwise loading in subsonic flow

NASA Technical Reports Server (NTRS)

Katzoff, S; Faison, M Frances; Dubose, Hugh C

1954-01-01

The field of a uniformly loaded wing in subsonic flow is discussed in terms of the acceleration potential. It is shown that, for the design of such wings, the slope of the mean camber surface at any point can be determined by a line integration around the wing boundary. By an additional line integration around the wing boundary, this method is extended to include the case where the local section lift coefficient varies with spanwise location (the chordwise loading at every section still remaining uniform). For the uniformly loaded wing of polygonal plan form, the integrations necessary to determine the local slope of the surface and the further integration of the slopes to determine the ordinate can be done analytically. An outline of these integrations and the resulting formulas are included. Calculated results are given for a sweptback wing with uniform chordwise loading and a highly tapered spanwise loading, a uniformly loaded delta wing, a uniformly loaded sweptback wing, and the same sweptback wing with uniform chordwise loading but elliptical span load distribution.

F-16XL ship #1 - CAWAP boundary layer rakes and hot film on left wing

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the boundary layer hot film and the boundary layer rakes on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 - CAWAP boundary layer hot film, left wing

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the boundary layer hot film on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. Hot film is used to measure temperature changes on a surface. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 CAWAP flight

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

An overview of the fundamental aerodynamics branch's research activities in wing leading-edge vortex flows at supersonic speeds

NASA Technical Reports Server (NTRS)

Miller, D. S.; Wood, R. M.; Covell, P. F.

1986-01-01

For the past 3 years, a research program pertaining to the study of wing leading edge vortices at supersonic speeds has been conducted in the Fundamental Aerodynamics Branch of the High-Speed Aerodynamics Division at the Langley Research Center. The purpose of the research is to provide an understanding of the factors governing the formation and the control of wing leading-edge vortices and to evaluate the use of these vortices for improving supersonic aerodynamic performance. The studies include both experimental and theoretical investigations and focus primarily on planform, thickness and camber effects for delta wings. An overview of this research activity is presented.

Bat breath reveals metabolic substrate use in free-ranging vampires.

PubMed

Voigt, Christian C; Grasse, Patricia; Rex, Katja; Hetz, Stefan K; Speakman, John R

2008-01-01

We analysed the stable carbon isotope ratio in exhaled CO(2) (delta(13)C(breath)) of free-ranging vampires to assess the type of metabolized substrate (endogenous or exogenous substrate) and its origin, i.e. whether the carbon atoms came from a C(4) food web (grass and cattle) or the C(3) food web in which they were captured (a rainforest remnant and its mammals). For an improved understanding of factors influencing the delta(13)C(breath) of vampires, we conducted feeding experiments with captive animals. The mean delta(13)C(breath) of starved bats was depleted in (13)C in relation to the diet by 4.6 per thousand (n = 10). Once fed with blood, delta(13)C(breath )levelled off within a short time approximately 2.2 per thousand above the stable carbon isotope signature of the diet. The median time required to exchange 50% of the carbon atoms in exhaled CO(2) with carbon atoms from the ingested blood was 18.6 min (mean 29.5 +/- 19.0 min, n = 5). The average delta(13)C of wing membrane and fur in free-ranging vampire bats suggested that bats almost exclusively foraged for cattle blood during the past weeks. The delta(13)C(breath) of the same bats averaged -19.1 per thousand. Given that all free-ranging vampires were starving and that the delta(13)C of cattle was more in enriched in (13)C by 5-6 per thousand than the delta(13)C(breath) of vampires, we conclude that the vampire bats of our study metabolised fat that was predominantly built from carbon atoms originating from cattle blood. Since delta(13)C of wing membrane and fur integrates over weeks and months respectively and delta(13)C(breath) over hours and days, we also conclude that vampire bats of the studied population consistently ignored rainforest mammals and chose cattle as their prey during and prior to our study.

Vapor-screen flow-visualization experiments in the NASA Langley 0.3-m transonic cryogenic tunnel

NASA Technical Reports Server (NTRS)

Selby, G. V.

1986-01-01

The vortical flow on the leeward side of a delta-wing model has been visualized at several different tunnel conditions in the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel using a vapor-screen flow-visualization technique. Vapor-screen photographs of the subject flow field are presented and interpreted relative to phenomenological implications. Results indicate that the use of nitrogen fog in conjunction with the vapor-screen technique is feasibile.

Detailed flow-field measurements over a 75 deg swept delta wing

NASA Technical Reports Server (NTRS)

Kjelgaard, Scott O.; Sellers, William L., III

1990-01-01

Results from an experimental investigation documenting the flowfield over a 75 deg swept delta wing at an angle-of-attack of 20.5 deg are presented. Results obtained include surface flow visualization, off-body flow visualization, and detailed flowfield surveys for various Reynolds numbers. Flowfield surveys at Reynolds numbers of 0.5, 1.0, and 1.5 million based on the root chord were conducted with both a Pitot pressure probe and a 5-hole pressure probe; and 3-component laser velocimeter surveys were conducted at a Reynolds number of 1.0 million. The Pitot pressure surveys were obtained at 5 chordwise stations, the 5-hole probe surveys were obtained at 3 chordwise stations and the laser velocimeter surveys were obtained at one station. The results confirm the classical roll up of the flow into a pair of primary vortices over the delta wing. The velocity measurements indicate that Reynolds number has little effect on the global structure of the flowfield for the Reynolds number range investigated. Measurements of the non-dimensional axial velocity in the core of the vortex indicate a jet like flow with values greater than twice freestream. Comparisons between velocity measurements from the 5-hole pressure probe and the laser velocimeter indicate that the pressure probe does a reasonable job of measuring the flowfield quantities where the velocity gradients in the flowfield are low.

Slender wing theory including regions of embedded total pressure loss

NASA Technical Reports Server (NTRS)

Mccune, James E.; Tavares, T. Sean; Lee, Norman K. W.; Weissbein, David

1988-01-01

An aerodynamic theory of the flow about slender delta wings is described. The theory includes a treatment of the self-consistent development of the vortex wake patterns above the wing necessary to maintain smooth flow at the wing edges. The paper focuses especially on the formation within the wake of vortex 'cores' as embedded regions of total pressure loss, fed and maintained by umbilical vortex sheets emanating from the wing edges. Criteria are developed for determining the growing size and location of these cores, as well as the distribution and strength of the vorticity within them. In this paper, however, the possibility of vortex breakup is omitted. The aerodynamic consequences of the presence and evolution of the cores and the associated wake structure are illustrated and discussed. It is noted that wake history effects can have substantial influence on the distribution of normal force on the wing as well as on its magnitude.

F-16XL ship #1 - CAWAP outboard rakes #7 and inboard rack #3

NASA Technical Reports Server (NTRS)

1996-01-01

This photo shows the #7 outboard rake and the #3 inboard rake on the left wing of NASA's single-seat F-16XL (ship #1) used for the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 CAWAP flight

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

76 FR 304 - Endangered and Threatened Wildlife and Plants; 90-Day Finding on a Petition To List the Red Knot...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-01-04

...). Juveniles resemble non-breeding adults, except that the feathers of the scapulars and wing coverts of... stopover sites are Grays Harbor and Willapa Bay in Washington, and Yukon-Kuskokwim Delta and Copper River Delta in Alaska (Isleib 1979, p. 128; Gill and Handel 1990, p. 712; Page et al. 1999, p. 467). Smaller...

75 FR 44094 - Update of the Federal Energy Regulatory Commission's Fees Schedule for Annual Charges for the Use...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-07-28

... 15.69 Colorado Crowley 7.85 Colorado Custer 47.08 Colorado Delta 62.78 Colorado Denver 31.39 Colorado... Michigan Delta 47.08 Michigan Dickinson 47.08 Michigan Eaton 94.17 Michigan Emmet 94.17 Michigan Genesee... Cottonwood......... 47.08 Minnesota Crow Wing 31.39 Minnesota Dakota 94.17 Minnesota Dodge 62.78 Minnesota...

Space shuttle: High angle of attack transition and low angle of attack launch phase aerodynamic stability and control of GD/C B-18E-2, B-18E-3 delta wing booster, and launch configuration of MSC-040A orbiter and twin pressure fed boosters

NASA Technical Reports Server (NTRS)

Debevoise, J. M.; Mcginnis, R. F.

1972-01-01

The test was a conventional stability and control test except for two aspects. One was the very high angles of attack at which the delta wing configurations were tested (up to 60 degrees) at Mach numbers of 3 and 4.96. The other was the installation of the orbiter and twin boosters in a manner that caused the support system to induce normal forces and side forces on the aft portion of the boosters at all Mach numbers; i.e., the support and the booster bodies were close together, side by side.

Flutter analysis of low aspect ratio wings

NASA Technical Reports Server (NTRS)

Parnell, L. A.

1986-01-01

Several very low aspect ratio flat plate wing configurations are analyzed for their aerodynamic instability (flutter) characteristics. All of the wings investigated are delta planforms with clipped tips, made of aluminum alloy plate and cantilevered from the supporting vehicle body. Results of both subsonic and supersonic NASTRAN aeroelastic analyses as well as those from another version of the program implementing the supersonic linearized aerodynamic theory are presented. Results are selectively compared with the experimental data; however, supersonic predictions of the Mach Box method in NASTRAN are found to be erratic and erroneous, requiring the use of a separate program.

Prediction of in-depth gap heating ratios from wing glove model test data. [space shuttle orbiter

NASA Technical Reports Server (NTRS)

1977-01-01

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.

24 CFR 3280.305 - Structural design requirements.

Code of Federal Regulations, 2010 CFR

2010-04-01

... Kandiyohi Carver Blue Earth Cass Meeker Dakota Martin Crow Wing Wright Goodhue Watonwan Aitkin Lac qui Parle... Charlevoix Benzie Marquette Menominee Montmorency Grand Traverse Alger Delta Alpena Kalkaska Luce Schoolcraft...

24 CFR 3280.305 - Structural design requirements.

Code of Federal Regulations, 2011 CFR

2011-04-01

... Kandiyohi Carver Blue Earth Cass Meeker Dakota Martin Crow Wing Wright Goodhue Watonwan Aitkin Lac qui Parle... Charlevoix Benzie Marquette Menominee Montmorency Grand Traverse Alger Delta Alpena Kalkaska Luce Schoolcraft...

Application of the CAP-TSD unsteady transonic small disturbance program to wing flutter. [Computational Aeroelasticity Program

NASA Technical Reports Server (NTRS)

Bennett, Robert M.; Batina, John T.

1989-01-01

The application and assessment of a computer program called CAP-TSD (Computational Aeroelasticity Program - Transonic Small Disturbance) for flutter predictions are described. Flutter calculations are presented for two thin swept-and-tapered wing planforms with well-defined modal properties. One planform is a series of 45-degree swept wings and the other planform is a clipped delta wing. Comparisons are made between the results of CAP-TSD using the linear equation and no airfoil thickness and the results obtained from a subsonic kernel function analysis. The calculations cover a Mach number range from low subsonic to low supersonic values, including the transonic range, and are compared with subsonic linear theory and experimental data. It is noted that since both wings have very thin airfoil sections, the effects of thickness are minimal.

A computer program for calculating aerodynamic characteristics of low aspect-ratio wings with partial leading-edge separation

NASA Technical Reports Server (NTRS)

Mehrotra, S. C.; Lan, C. E.

1978-01-01

The necessary information for using a computer program to predict distributed and total aerodynamic characteristics for low aspect ratio wings with partial leading-edge separation is presented. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the Quasi-Vortex-Lattice method. The leading edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at midpoints to satisfy the force free condition. The wake behind the trailing edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading and trailing edges. The program is restricted to delta wings with zero thickness and no camber. It is written in FORTRAN language and runs on CDC 6600 computer.

Eclipse program QF-106 aircraft in flight

NASA Technical Reports Server (NTRS)

1997-01-01

This photo shows one of the QF-106s used in the Eclipse project in flight. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Topological structures of vortex flow on a flying wing aircraft, controlled by a nanosecond pulse discharge plasma actuator

NASA Astrophysics Data System (ADS)

Du, Hai; Shi, Zhiwei; Cheng, Keming; Wei, Dechen; Li, Zheng; Zhou, Danjie; He, Haibo; Yao, Junkai; He, Chengjun

2016-06-01

Vortex control is a thriving research area, particularly in relation to flying wing or delta wing aircraft. This paper presents the topological structures of vortex flow on a flying wing aircraft controlled by a nanosecond plasma dielectric barrier discharge actuator. Experiments, including oil flow visualization and two-dimensional particle image velocimetry (PIV), were conducted in a wind tunnel with a Reynolds number of 0.5 × 106. Both oil and PIV results show that the vortex can be controlled. Oil topological structures on the aircraft surface coincide with spatial PIV flow structures. Both indicate vortex convergence and enhancement when the plasma discharge is switched on, leading to a reduced region of separated flow.

On vortex bursting

NASA Technical Reports Server (NTRS)

Werle, H.

1984-01-01

Vortex bursting is studied by means of visualization. The physical behavior of the phenomenon is emphasized, and its similarity with boundary layer separation or wake bursting becomes apparent. The essential influence of an increasing pressure gradient on the initiation, the position and the type of bursting is clearly confirmed. The evolution of the phenomena as a function of several parameters is analyzed in the case of delta wings, alone or installed on aircraft models, and compared with the results of similar wind tunnel or flight tests.

Space shuttle: Heat transfer rate measurements on Convair booster (B-15B-2) at nominal Mach number of 8

NASA Technical Reports Server (NTRS)

Warmbrod, J. D.; Martindale, W. R.; Matthews, R. K.

1971-01-01

Plotted and tabulated data on heat transfer from a thin-skin thermocouple are presented. The data is representative of the reentry event of the booster alone configuration. The data were generated during wind tunnel tests of the B-15B-2 delta wing booster at Mach 8. Thermocouple measurements are reduced to heat transfer coefficient ratio and the data are presented as plotted variations versus longitudinal, lateral, and vertical local model positions.

7 CFR 1000.52 - Adjusted Class I differentials.

Code of Federal Regulations, 2011 CFR

2011-01-01

....45 CONEJOS CO 08021 1.90 COSTILLA CO 08023 1.90 CROWLEY CO 08025 2.45 CUSTER CO 08027 2.45 DELTA CO... 1.80 CRAWFORD MI 26039 1.80 DELTA MI 26041 1.70 DICKINSON MI 26043 1.70 EATON MI 26045 1.80 EMMET MI... MN 27029 1.65 COOK MN 27031 1.65 COTTONWOOD MN 27033 1.70 CROW WING MN 27035 1.65 DAKOTA MN 27037 1...

7 CFR 1000.52 - Adjusted Class I differentials.

Code of Federal Regulations, 2010 CFR

2010-01-01

....45 CONEJOS CO 08021 1.90 COSTILLA CO 08023 1.90 CROWLEY CO 08025 2.45 CUSTER CO 08027 2.45 DELTA CO... 1.80 CRAWFORD MI 26039 1.80 DELTA MI 26041 1.70 DICKINSON MI 26043 1.70 EATON MI 26045 1.80 EMMET MI... MN 27029 1.65 COOK MN 27031 1.65 COTTONWOOD MN 27033 1.70 CROW WING MN 27035 1.65 DAKOTA MN 27037 1...

KSC-98pc1191

NASA Image and Video Library

1998-09-30

KENNEDY SPACE CENTER, FLA. -- Deep Space 1 is lifted from its work platform, giving a closer view of the experimental solar-powered ion propulsion engine. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. Above the engine is one of the two solar wings, folded for launch, that will provide the power for it. When fully extended, the wings measure 38.6 feet from tip to tip. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Another onboard experiment includes software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches

KSC-98pc1189

NASA Image and Video Library

1998-09-30

KENNEDY SPACE CENTER, FLA. -- Deep Space 1 rests on its work platform after being fitted with thermal insulation. The reflective insulation is designed to protect the spacecraft as this side faces the sun. At either side of the spacecraft are its solar wings, folded for launch. When fully extended, the wings measure 38.6 feet from tip to tip. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include a solar-powered ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches

Deep Space 1 moves to CCAS for testing

NASA Technical Reports Server (NTRS)

1998-01-01

Workers in the Payload Hazardous Servicing Facility lower Deep Space 1 onto its transporter, for movement to the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station, where it will undergo testing. At either side of the spacecraft are its solar wings, folded for launch. When fully extended, the wings measure 38.6 feet from tip to tip. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include a solar-powered ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches.

KSC-98pc1188

NASA Image and Video Library

1998-09-30

KENNEDY SPACE CENTER, FLA. -- Workers in the Payload Hazardous Servicing Facility lower Deep Space 1 onto its transporter, for movement to the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station, where it will undergo testing. At either side of the spacecraft are its solar wings, folded for launch. When fully extended, the wings measure 38.6 feet from tip to tip. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include a solar-powered ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches

KSC-98pc1190

NASA Image and Video Library

1998-09-30

KENNEDY SPACE CENTER, FLA. -- Deep Space 1 rests on its work platform after being fitted with thermal insulation. The dark insulation is designed to protect the side of the spacecraft that faces away from the sun. At either side of the spacecraft are its solar wings, folded for launch. When fully extended, the wings measure 38.6 feet from tip to tip. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include a solar-powered ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches

23 CFR Appendix A to Part 658 - National Network-Federally-Designated Routes

Code of Federal Regulations, 2011 CFR

2011-04-01

... Anchorage AK 3 Palmer. AK 2 AK 3 Fairbanks Milepost 1412 Delta Junction. AK 3 AK 1 Palmer AK 2 Fairbanks... Harbors. US 63 I-90 Rochester US 52 Rochester. US 63 MN 58 Red Wing WI State Line. US 71 IA State Line MN... State Line Red Wing MN US 2 W. of Ashland. US 141 US 41 Abrams US 8 Pembine. US 151 IA State Line...

23 CFR Appendix A to Part 658 - National Network-Federally-Designated Routes

Code of Federal Regulations, 2010 CFR

2010-04-01

... Anchorage AK 3 Palmer. AK 2 AK 3 Fairbanks Milepost 1412 Delta Junction. AK 3 AK 1 Palmer AK 2 Fairbanks... Harbors. US 63 I-90 Rochester US 52 Rochester. US 63 MN 58 Red Wing WI State Line. US 71 IA State Line MN... State Line Red Wing MN US 2 W. of Ashland. US 141 US 41 Abrams US 8 Pembine. US 151 IA State Line...

Control of Interacting Vortex Flows at Subsonic and Transonic Speeds Using Passive Porosity

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2003-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) 8-Foot Transonic Pressure Tunnel (TPT) to determine the effects of passive surface porosity on vortex flow interactions about a general research fighter configuration at subsonic and transonic speeds. Flow-through porosity was applied to a wing leading-edge extension (LEX) mounted to a 65 deg cropped delta wing model to promote large nose-down pitching moment increments at high angles of attack. Porosity decreased the vorticity shed from the LEX, which weakened the LEX vortex and altered the global interactions of the LEX and wing vortices at high angles of attack. Six-component forces and moments and wing upper surface static pressure distributions were obtained at free-stream Mach numbers of 0.50, 0.85, and 1.20, Reynolds number of 2.5(10(exp 6)) per foot, angles of attack up to 30 deg, and angles of sideslip to +/- 8 deg. The off-surface flow field was visualized in selected cross-planes using a laser vapor screen flow visualization technique. Test data were obtained with a centerline vertical tail and with alternate twin, wing-mounted vertical fins having 0 deg and 30 deg cant angles. In addition, the porosity of the LEX was compartmentalized to determine the sensitivity of the vortex-dominated aerodynamics to the location and level of porosity applied to the LEX.

The formation mechanism and impact of streamwise vortices on NACA 0021 airfoil's performance with undulating leading edge modification

NASA Astrophysics Data System (ADS)

Rostamzadeh, N.; Hansen, K. L.; Kelso, R. M.; Dally, B. B.

2014-10-01

Wings with tubercles have been shown to display advantageous loading behavior at high attack angles compared to their unmodified counterparts. In an earlier study by the authors, it was shown that an undulating leading-edge configuration, including but not limited to a tubercled model, induces a cyclic variation in circulation along the span that gives rise to the formation of counter-rotating streamwise vortices. While the aerodynamic benefits of full-span tubercled wings have been associated with the presence of such vortices, their formation mechanism and influence on wing performance are still in question. In the present work, experimental and numerical tests were conducted to further investigate the effect of tubercles on the flow structure over full-span modified wings based on the NACA 0021 profile, in the transitional flow regime. It is found that a skew-induced mechanism accounts for the formation of streamwise vortices whose development is accompanied by flow separation in delta-shaped regions near the trailing edge. The presence of vortices is detrimental to the performance of full-span wings pre-stall, however renders benefits post-stall as demonstrated by wind tunnel pressure measurement tests. Finally, primary and secondary vortices are identified post-stall that produce an enhanced momentum transfer effect that reduces flow separation, thus increasing the generated amount of lift.

Supersonic aerodynamic characteristics of the North American Rockwell ATP shuttle orbiter

NASA Technical Reports Server (NTRS)

Ware, G. M.; Pencer, B., Jr.; Founier, R. H.

1973-01-01

A wind tunnel study to determine the supersonic aerodynamic characteristics of a 0.01925-scale model of the space shuttle orbiter configuration is reported. The model consisted of a low-finess-ratio body with a blended 50 swept delta wing forming an ogee planform and a center-line-mounted vertical tail. Tests were made at Mach numbers from 1.90 to 4.63, at angles of attack from -6 to 30, at angles of sideslip of 0 and 3, and at a Reynolds number, based on body length, of 5.3x 1 million.

Lift Augmentation on a Delta Wing via Leading Edge Fences and the Gurney Flap

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.; Tso, Jin

1993-01-01

Wind tunnel tests have been conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve by as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

Experimental Investigation of the Flow Structure over a Delta Wing Via Flow Visualization Methods.

PubMed

Shen, Lu; Chen, Zong-Nan; Wen, Chihyung

2018-04-23

It is well known that the flow field over a delta wing is dominated by a pair of counter rotating leading edge vortices (LEV). However, their mechanism is not well understood. The flow visualization technique is a promising non-intrusive method to illustrate the complex flow field spatially and temporally. A basic flow visualization setup consists of a high-powered laser and optic lenses to generate the laser sheet, a camera, a tracer particle generator, and a data processor. The wind tunnel setup, the specifications of devices involved, and the corresponding parameter settings are dependent on the flow features to be obtained. Normal smoke wire flow visualization uses a smoke wire to demonstrate the flow streaklines. However, the performance of this method is limited by poor spatial resolution when it is conducted in a complex flow field. Therefore, an improved smoke flow visualization technique has been developed. This technique illustrates the large-scale global LEV flow field and the small-scale shear layer flow structure at the same time, providing a valuable reference for later detailed particle image velocimetry (PIV) measurement. In this paper, the application of the improved smoke flow visualization and PIV measurement to study the unsteady flow phenomena over a delta wing is demonstrated. The procedure and cautions for conducting the experiment are listed, including wind tunnel setup, data acquisition, and data processing. The representative results show that these two flow visualization methods are effective techniques for investigating the three-dimensional flow field qualitatively and quantitatively.

Controlled vortical flow on delta wings through unsteady leading edge blowing

NASA Technical Reports Server (NTRS)

Lee, K. T.; Roberts, Leonard

1990-01-01

The vortical flow over a delta wing contributes an important part of the lift - the so called nonlinear lift. Controlling this vortical flow with its favorable influence would enhance aircraft maneuverability at high angle of attack. Several previous studies have shown that control of the vortical flow field is possible through the use of blowing jets. The present experimental research studies vortical flow control by applying a new blowing scheme to the rounded leading edge of a delta wing; this blowing scheme is called Tangential Leading Edge Blowing (TLEB). Vortical flow response both to steady blowing and to unsteady blowing is investigated. It is found that TLEB can redevelop stable, strong vortices even in the post-stall angle of attack regime. Analysis of the steady data shows that the effect of leading edge blowing can be interpreted as an effective change in angle of attack. The examination of the fundamental time scales for vortical flow re-organization after the application of blowing for different initial states of the flow field is studied. Different time scales for flow re-organization are shown to depend upon the effective angle of attack. A faster response time can be achieved at angles of attack beyond stall by a suitable choice of the initial blowing momentum strength. Consequently, TLEB shows the potential of controlling the vortical flow over a wide range of angles of attack; i.e., in both for pre-stall and post-stall conditions.

High transonic speed transport aircraft study. [aerodynamic characteristics of single-fuselage, yawed-wing configuration

NASA Technical Reports Server (NTRS)

Kulfan, R. M.; Neumann, F. D.; Nisbet, J. W.; Mulally, A. R.; Murakami, J. K.; Noble, E. C.; Mcbarron, J. P.; Stalter, J. L.; Gimmestad, D. W.; Sussman, M. B.

1973-01-01

An initial design study of high-transonic-speed transport aircraft has been completed. Five different design concepts were developed. These included fixed swept wing, variable-sweep wing, delta wing, double-fuselage yawed-wing, and single-fuselage yawed-wing aircraft. The boomless supersonic design objectives of range=5560 Km (3000 nmi), payload-18 143 kg (40 000lb), Mach=1.2, and FAR Part 36 aircraft noise levels were achieved by the single-fuselage yawed-wing configuration with a gross weight of 211 828 Kg (467 000 lb). A noise level of 15 EPNdB below FAR Part 36 requirements was obtained with a gross weight increase to 226 796 Kg (500 000 lb). Although wing aeroelastic divergence was a primary design consideration for the yawed-wing concepts, the graphite-epoxy wings of this study were designed by critical gust and maneuver loads rather than by divergence requirements. The transonic nacelle drag is shown to be very sensitive to the nacelle installation. A six-degree-of-freedom dynamic stability analysis indicated that the control coordination and stability augmentation system would require more development than for a symmetrical airplane but is entirely feasible. A three-phase development plan is recommended to establish the full potential of the yawed-wing concept.

F-16XL ship #1 CAWAP flight - alpha 5 degrees, altitude 10,000 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 10,000 feet, with an angle of attack of 5 degrees. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

A Summary of Numerous Strain-Gage Load Calibrations on Aircraft Wings and Tails in a Technological Format

NASA Technical Reports Server (NTRS)

Jenkins, Jerald M.; DeAngelis, V. Michael

1997-01-01

Fifteen aircraft structures that were calibrated for flight loads using strain gages are examined. The primary purpose of this paper is to document important examples of load calibrations on airplanes during the past four decades. The emphasis is placed on studying the physical procedures of calibrating strain-gaged structures and all the supporting analyses and computational techniques that have been used. The results and experiences obtained from actual data from 14 structures (on 13 airplanes and 1 laboratory test structure) are presented. This group of structures includes fins, tails, and wings with a wide variety of aspect ratios. Straight- wing, swept-wing, and delta-wing configurations are studied. Some of the structures have skin-dominant construction; others are spar-dominant. Anisotropic materials, heat shields, corrugated components, nonorthogonal primary structures, and truss-type structures are particular characteristics that are included.

F-16XL ship #1 CAWAP flight - alpha 15 degrees, altitude 5,000 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 5000 feet, with an angle of attack of 15 degrees. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

78 FR 15602 - Annual Update to Fee Schedule for the Use of Government Lands by Hydropower Licensees

Federal Register 2010, 2011, 2012, 2013, 2014

2013-03-12

... Clear Creek 23.25 Conejos 23.14 Costilla 15.09 Crowley 9.35 Custer 30.38 Delta 53.59 Denver 19.36....40 Clinton 97.56 Crawford 109.40 Delta 61.65 Dickinson 66.61 Eaton 82.18 Emmet 96.58 Genesee 100.71... 74.17 Chisago 119.37 Clay 51.21 Clearwater 41.18 Cook 119.01 Cottonwood 82.17 Crow Wing 71.84 Dakota...

76 FR 53535 - Migratory Bird Hunting; Proposed Frameworks for Late-Season Migratory Bird Hunting Regulations

Federal Register 2010, 2011, 2012, 2013, 2014

2011-08-26

... habitats. These surveys are conducted using fixed- wing aircraft, helicopters, and ground crews and... Delta Waterfowl Foundation, the Max McGraw Wildlife Foundation, the LaCrosse County Conservation...

Detection probability of cliff-nesting raptors during helicopter and fixed-wing aircraft surveys in western Alaska

USGS Publications Warehouse

Booms, T.L.; Schempf, P.F.; McCaffery, B.J.; Lindberg, M.S.; Fuller, M.R.

2010-01-01

We conducted repeated aerial surveys for breeding cliff-nesting raptors on the Yukon Delta National Wildlife Refuge (YDNWR) in western Alaska to estimate detection probabilities of Gyrfalcons (Falco rusticolus), Golden Eagles (Aquila chrysaetos), Rough-legged Hawks (Buteo lagopus), and also Common Ravens (Corvus corax). Using the program PRESENCE, we modeled detection histories of each species based on single species occupancy modeling. We used different observers during four helicopter replicate surveys in the Kilbuck Mountains and five fixed-wing replicate surveys in the Ingakslugwat Hills near Bethel, AK. During helicopter surveys, Gyrfalcons had the highest detection probability estimate (p^;p^ 0.79; SE 0.05), followed by Golden Eagles (p^=0.68; SE 0.05), Common Ravens (p^=0.45; SE 0.17), and Rough-legged Hawks (p^=0.10; SE 0.11). Detection probabilities from fixed-wing aircraft in the Ingakslugwat Hills were similar to those from the helicopter in the Kilbuck Mountains for Gyrfalcons and Golden Eagles, but were higher for Common Ravens (p^=0.85; SE 0.06) and Rough-legged Hawks (p^=0.42; SE 0.07). Fixed-wing aircraft provided detection probability estimates and SEs in the Ingakslugwat Hills similar to or better than those from helicopter surveys in the Kilbucks and should be considered for future cliff-nesting raptor surveys where safe, low-altitude flight is possible. Overall, detection probability varied by observer experience and in some cases, by study area/aircraft type.

Experimental evaluation of nacelle-airframe interference forces and pressures at Mach numbers of 0.9 to 1.4

NASA Technical Reports Server (NTRS)

Bencze, D. P.

1977-01-01

Detailed interference force-and-pressure data were obtained on a representative supersonic transport wing-body-nacelle combination at Mach numbers of 0.9 to 1.4. The basic model consisted of a delta wing-body aerodynamic model with a length of 158.0 cm (62.2 in.) and a wingspan of 103.6 cm (40.8 in.) and four independently supported nacelles positioned beneath the model. The experimental program was conducted in the Ames 11- by 11-Foot Wind Tunnel at a constant unit Reynolds number. The primary variables examined included Mach number, angle of attack, nacelle position, and nacelle mass-flow ratio. Under the most favorable conditions, the net interference drag was equal to 50 percent the drag of four isolated nacelles at M = 1.4, 75 percent at M = 1.15, and 144 percent at M = 0.90. The overall interference effects were found to be rather constant over the operating angle-of-attack range of the configuration. The effects of mass-flow ratio on the interference pressure distributions were limited to the lip region of the nacelle and the local wing surface in the immediate vicinity of the nacelle lip. The net change in the measured interference forces resulting from variations in the nacelle mass-flow ratio were found to be quite small.

F-106 tow cable attachment and release mechanism for Eclipse program

NASA Technical Reports Server (NTRS)

1997-01-01

View of the tow cable attachment and release mechanism forward of the cockpit on the QF-106 Eclipse aircraft. This mechanism held and then released the Vectran rope used to tow the QF-106 behind an Air Force C-141A. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Closeup of QF-106 release hook for Eclipse program

NASA Technical Reports Server (NTRS)

1997-01-01

View of the release hook on the QF-106 that allowed the pilot to release the tow rope extending from the C-141A tow plane in the Eclipse project. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Eclipse program C-141A aircraft

NASA Technical Reports Server (NTRS)

1997-01-01

This photograph shows the Air Force C-141A that was used in the Eclipse project as a tow vehicle. The project used a QF-106 interceptor aircraft to simulate a future orbiter, which would be towed to a high altitude and released to fire its own engines and carry a payload into space. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

77 FR 49867 - Migratory Bird Hunting; Proposed Frameworks for Late-Season Migratory Bird Hunting Regulations

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-17

... fixed- wing aircraft, helicopters, and ground crews and encompass principal breeding areas of North.... Conditions throughout Alaska and northwestern Canada were good. The exception was the Yukon-Kuskokwim Delta...

Lift augmentation on a delta wing via leading edge fences and the Gurney flap. M.S. Thesis

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.

1992-01-01

Wind tunnel tests were conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

The design of two stage to orbit vehicles

NASA Astrophysics Data System (ADS)

Gregorek, G. M.; Ramsay, T. N.

1991-09-01

Two designs are presented for a two-stage-to-orbit vehicle to complement an existing heavy lift vehicle. The payload is 10,000 lbs and 27 ft long by 10 ft in diameter for design purposes and must be carried to a low earth orbit by an air-breathing carrier configuration that can take off horizontally within 15,000 ft. Two designs are presented: a delta wing/body carrier in which the fuselage contains the orbiter; and a cranked-delta wing/body carrier in which the orbiter is carried piggy back. The engines for both carriers are turbofanramjets powered with liquid hydrogen, and the orbiters employ either a Space Shuttle Main Engine or a half-scale version with additional scramjet engines. The orbiter based on a full-scale Space Shuttle Main Engine is found to have a significantly higher takeoff weight which results in a higher total takeoff weight.

Structural dynamics division research and technology accomplishments for fiscal year 1990 and plans for fiscal year 1991

NASA Technical Reports Server (NTRS)

Wynne, Eleanor C.

1991-01-01

The research accomplishments of the Structural Dynamics Division for F.Y. 1991 are presented. The work is discussed in terms of highlights of accomplishments during the past year and plans for the current year as they relate to 5-year plans and the objectives of each technical area. Included is research on unsteady aerodynamics, helicopter rotors, computational fluid dynamics, oscillations of leading edge flaps of a delta wing, and aircraft wing loads.

Aeroelastic, CFD, and Dynamics Computation and Optimization for Buffet and Flutter Applications

NASA Technical Reports Server (NTRS)

Kandil, Osama A.

1997-01-01

Accomplishments achieved during the reporting period are listed. These accomplishments included 6 papers published in various journals or presented at various conferences; 1 abstract submitted to a technical conference; production of 2 animated movies; and a proposal for use of the National Aerodynamic Simulation Facility at NASA Ames Research Center for further research. The published and presented papers and animated movies addressed the following topics: aeroelasticity, computational fluid dynamics, structural dynamics, wing and tail buffet, vortical flow interactions, and delta wings.

An Investigation of the Drag Characteristics of a Tailless Delta-Wing Airplane in Flight, Including Comparison with Wind-Tunnel Data

NASA Technical Reports Server (NTRS)

Rolls, L. Stewart; Wingrove, Rodney C.

1958-01-01

A series of flight tests were conducted to determine the lift and drag characteristics of an F4D-1 airplane over a Mach number range of 0.80 to 1.10 at an altitude of 40,000 feet. Apparently satisfactory agreement was obtained between the flight data and results from wind-tunnel tests of an 0.055-scale model of the airplane. Further tests show the apparent agreement was a consequence of the altitude at which the first tests were made.

Heat transfer distributions on the LMSC 040C and 040A-L4 delta wing orbiters (M equals 8)7

NASA Technical Reports Server (NTRS)

Baker, R. C.; Mcgee, K. W.; Schultz, H. D.

1972-01-01

The results of a wind tunnel investigation are presented for measuring aerodynamic heat transfer distributions on the 040C and 040A-L4 space shuttle orbiter configurations. Heat transfer rates were determined by the phase change coating technique, using 0.012-scale Stycast models coated with Tempilaq. Data were obtained at a nominal free stream Mach number of 8, Reynolds numbers from 1.0 to 3.8 million per foot, and angles of attack from 20 to 60 deg.

LAVA Simulations for the AIAA Sonic Boom Prediction Workshop

NASA Technical Reports Server (NTRS)

Housman, Jeffrey A.; Sozer, Emre; Moini-Yekta , Shayan; Kiris, Cetin C.

2014-01-01

Computational simulations using the Launch Ascent and Vehicle Aerodynamics (LAVA) framework are presented for the First AIAA Sonic Boom Prediction Workshop test cases. The framework is utilized with both structured overset and unstructured meshing approaches. The three workshop test cases include an axisymmetric body, a Delta Wing-Body model, and a complete low-boom supersonic transport concept. Solution sensitivity to mesh type and sizing, and several numerical convective flux discretization choices are presented and discussed. Favorable comparison between the computational simulations and experimental data of nearand mid-field pressure signatures were obtained.

Drosophila Lyra mutations are gain-of-function mutations of senseless

NASA Technical Reports Server (NTRS)

Nolo, R.; Abbott, L. A.; Bellen, H. J.

2001-01-01

The Lyra mutation was first described by Jerry Coyne in 1935. Lyra causes recessive pupal lethality and adult heterozygous Lyra mutants exhibit a dominant loss of the anterior and posterior wing margins. Unlike many mutations that cause loss of wing tissue (e.g., scalloped, Beadex, cut, and apterous-Xasta), Lyra wing discs do not exhibit increased necrotic or apoptotic cell death, nor do they show altered BrdU incorporation. However, during wing disc eversion, loss of the anterior and posterior wing margins is apparent. We have previously shown that senseless, a gene that is necessary and sufficient for peripheral nervous system (PNS) development, is allelic to Lyra. Here we show by several genetic criteria that Lyra alleles are neomorphic alleles of senseless that cause ectopic expression of SENSELESS in the wing pouch. Similarly, overexpression of SENSELESS in the wing disc causes loss of wing margin tissue, thereby mimicking the Lyra phenotype. Lyra mutants display aberrant expression of DELTA, VESTIGIAL, WINGLESS, and CUT. As in Lyra mutants, overexpression of SENSELESS in some areas of the wing pouch also leads to loss of WINGLESS and CUT. In summary, our data indicate that overexpression of SENSELESS causes a severe reduction in NOTCH signaling that in turn may lead to decreased transcription of several key genes required for wing development, leading to a failure in cell proliferation and loss of wing margin tissue.

An efficient coordinate transformation technique for unsteady, transonic aerodynamic analysis of low aspect-ratio wings

NASA Technical Reports Server (NTRS)

Guruswamy, G. P.; Goorjian, P. M.

1984-01-01

An efficient coordinate transformation technique is presented for constructing grids for unsteady, transonic aerodynamic computations for delta-type wings. The original shearing transformation yielded computations that were numerically unstable and this paper discusses the sources of those instabilities. The new shearing transformation yields computations that are stable, fast, and accurate. Comparisons of those two methods are shown for the flow over the F5 wing that demonstrate the new stability. Also, comparisons are made with experimental data that demonstrate the accuracy of the new method. The computations were made by using a time-accurate, finite-difference, alternating-direction-implicit (ADI) algorithm for the transonic small-disturbance potential equation.

Prediction and control of slender-wing rock

NASA Technical Reports Server (NTRS)

Kandil, Osama A.; Salman, Ahmed A.

1992-01-01

The unsteady Euler equations and the Euler equations of rigid-body dynamics, both written in the moving frame of reference, are sequentially solved to simulate the limit-cycle rock motion of slender delta wings. The governing equations of the fluid flow and the dynamics of the present multidisciplinary problem are solved using an implicit, approximately-factored, central-difference-like, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. For the control of wing-rock motion, leading-edge flaps are forced to oscillate anti-symmetrically at prescribed frequency and amplitude, which are tuned in order to suppress the rock motion. Since the computational grid deforms due to the leading-edge flaps motion, the grid is dynamically deformed using the Navier-displacement equations. Computational applications cover locally-conical and three-dimensional solutions for the wing-rock simulation and its control.

Space shuttle: Aerodynamic heating tests of the MDAC delta wing orbiter and canard booster

NASA Technical Reports Server (NTRS)

Andresen, T. L.

1972-01-01

Design of an efficient thermal protection system for the space shuttle orbiter and booster is discussed, based on knowledge of the thermal environment to be experienced by the vehicles in all flight phases. The complex configurations of these vehicles limit the level of confidence which can be associated with purely analytical thermal environment predictions. Tests were conducted during April and May 1971 using an orbiter and booster model at a 96-in. hypersonic shock tunnel. Both models were tested separately as well as together. A sufficiently large range in Reynolds number was covered so that laminar, transitional, and turbulent data could be obtained.

Heat Transfer and Fluid Mechanics Institute, 24th, Oregon State University, Corvallis, Ore., June 12-14, 1974, Proceedings

NASA Technical Reports Server (NTRS)

Davis, L. R. (Editor); Wilson, R. E.

1974-01-01

Recent theoretical and experimental studies in heat transfer and fluid mechanics, including some environmental protection investigations, are presented in a number of papers. Some of the topics covered include condensation heat transfer, a model of turbulent momentum and heat transfer at points of separation and reattachment, an explicit scheme for calculations of confined turbulent flows with heat transfer, heat transfer effects on a delta wing in subsonic flow, fluid mechanics of ocean outfalls, thermal plumes from industrial cooling water, a photochemical air pollution model for the Los Angeles air basin, and a turbulence model of diurnal variations in the planetary boundary layer. Individual items are announced in this issue.

78 FR 52337 - Migratory Bird Hunting; Proposed Frameworks for Late-Season Migratory Bird Hunting Regulations

Federal Register 2010, 2011, 2012, 2013, 2014

2013-08-22

... conducted using fixed- wing aircraft, helicopters, and ground crews and encompass principal breeding areas... conditions in the north-country was Alaska's Yukon-Kuskokwim Delta (YKD), where ice break-up was the latest...

Multidisciplinary Shape Optimization of a Composite Blended Wing Body Aircraft

NASA Astrophysics Data System (ADS)

Boozer, Charles Maxwell

A multidisciplinary shape optimization tool coupling aerodynamics, structure, and performance was developed for battery powered aircraft. Utilizing high-fidelity computational fluid dynamics analysis tools and a structural wing weight tool, coupled based on the multidisciplinary feasible optimization architecture; aircraft geometry is modified in the optimization of the aircraft's range or endurance. The developed tool is applied to three geometries: a hybrid blended wing body, delta wing UAS, the ONERA M6 wing, and a modified ONERA M6 wing. First, the optimization problem is presented with the objective function, constraints, and design vector. Next, the tool's architecture and the analysis tools that are utilized are described. Finally, various optimizations are described and their results analyzed for all test subjects. Results show that less computationally expensive inviscid optimizations yield positive performance improvements using planform, airfoil, and three-dimensional degrees of freedom. From the results obtained through a series of optimizations, it is concluded that the newly developed tool is both effective at improving performance and serves as a platform ready to receive additional performance modules, further improving its computational design support potential.

F-16XL ship #1 CAWAP flight - alpha 10 degrees, beta -5 degrees, altitude 10,000 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 10,000 feet, with an angle of attack of 10 degrees and a sideslip angle of -5 degrees. The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

F-16XL ship #1 CAWAP flight - alpha 21 degrees, altitude 17,500 feet

NASA Technical Reports Server (NTRS)

1996-01-01

The single-seat F-16XL (ship #1) makes another run during the Cranked-Arrow Wing Aerodynamic Project (CAWAP) at Dryden Flight Research Center, Edwards, California. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing (visible here) has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. This photo shows the aircraft gathering data at an altitude of 17,500 feet, with an angle of attack of 21 degrees The program also gathered aero data on two wing planforms for NASA's High Speed Research Program. The first flight of CAWAP occurred on November 21, 1995, and the test program ended in April 1996.

An aerodynamic assessment of various supersonic fighter airplanes based on Soviet design concepts

NASA Technical Reports Server (NTRS)

Spearman, M. L.

1983-01-01

The aerodynamic, stability, and control characteristics of several supersonic fighter airplane concepts were assessed. The configurations include fixed-wing airplanes having delta wings, swept wings, and trapezoidal wings, and variable wing-sweep airplanes. Each concept employs aft tail controls. The concepts vary from lightweight, single engine, air superiority, point interceptor, or ground attack types to larger twin-engine interceptor and reconnaissance designs. Results indicate that careful application of the transonic or supersonic area rule can provide nearly optimum shaping for minimum drag for a specified Mach number requirement. Through the proper location of components and the exploitation of interference flow fields, the concepts provide linear pitching moment characteristics, high control effectiveness, and reasonably small variations in aerodynamic center location with a resulting high potential for maneuvering capability. By careful attention to component shaping and location and through the exploitation of local flow fields, favorable roll-to-yaw ratios may result and a high degree of directional stability can be achieved.

A quantitative comparison of leading-edge vortices in incompressible and supersonic flows

DOT National Transportation Integrated Search

2002-01-14

When requiring quantitative data on delta-wing vortices for design purposes, low-speed results have often been extrapolated to configurations intended for supersonic operation. This practice stems from a lack of database owing to difficulties that pl...

Active Flow Separation Control on a NACA 0015 Wing Using Fluidic Actuators

NASA Technical Reports Server (NTRS)

Melton, Latunia P.

2014-01-01

Results are presented from a recent set of wind tunnel experiments using sweeping jet actuators to control ow separation on the 30% chord trailing edge ap of a 30 deg. swept wing model with an aspect ratio (AR) of 4.35. Two sweeping jet actuator locations were examined, one on the flap shoulder and one on the trailing edge flap. The parameters that were varied included actuator momentum, freestream velocity, and trailing edge flap deflection (Delta f ) angle. The primary focus of this set of experiments was to determine the mass flow and momentum requirements for controlling separation on the flap, especially at large flap deflection angles which would be characteristic of a high lift system. Surface pressure data, force and moment data, and stereoscopic particle image velocimetry (PIV) data were acquired to evaluate the performance benefits due to applying active flow control. Improvements in lift over the majority of the wing span were obtained using sweeping jet actuator control. High momentum coefficient, Cu, levels were needed when using the actuators on the ap because they were located downstream of separation. Actuators on the flap shoulder performed slightly better but actuator size, orientation, and spacing still need to be optimized.

Hang Gliders

NASA Technical Reports Server (NTRS)

1981-01-01

Francis M. Rogallo and his wife Gertrude researched flexible controllable fabric airfoils with a delta, V-shaped, configuration for use on inexpensive private aircraft. They were issued a flex-wing patent and refined their designs. Development of Rogallo wings, used by U.S. Moyes, Inc. substantially broadened the flexible airfoil technology base which originated from NASA's reentry parachute. The Rogallo technology, particularly the airfoil frame was incorporated in the design of a kite by John Dickenson. The Dickenson kite served as prototype for the Australian Moyes line of hang gliders. Company no longer exists.

KSC-98pc1187

NASA Image and Video Library

1998-09-30

KENNEDY SPACE CENTER, FLA. -- After covering the bulk of Deep Space 1 in thermal insulating blankets, workers in the Payload Hazardous Servicing Facility lift it from its work platform before moving it onto its transporter (behind workers at left). Deep Space 1 is being moved to the Defense Satellite Communications System Processing Facility (DPF), Cape Canaveral Air Station, for testing. At either side of the spacecraft are its solar wings, folded for launch. When fully extended, the wings measure 38.6 feet from tip to tip. The first flight in NASA's New Millennium Program, Deep Space 1 is designed to validate 12 new technologies for scientific space missions of the next century. Onboard experiments include a solar-powered ion propulsion engine and software that tracks celestial bodies so the spacecraft can make its own navigation decisions without the intervention of ground controllers. The ion propulsion engine is the first non-chemical propulsion to be used as the primary means of propelling a spacecraft. Deep Space 1 will complete most of its mission objectives within the first two months, but may also do a flyby of a near-Earth asteroid, 1992 KD, in July 1999. Deep Space 1 will be launched aboard a Boeing Delta 7326 rocket from Launch Pad 17A, Cape Canaveral Air Station, in October. Delta II rockets are medium capacity expendable launch vehicles derived from the Delta family of rockets built and launched since 1960. Since then there have been more than 245 Delta launches

75 FR 62810 - Twenty-Fourth Update of the Federal Agency Hazardous Waste Compliance Docket

Federal Register 2010, 2011, 2012, 2013, 2014

2010-10-13

... Alliance CCAFS Beach Road CCAFS FL 32920-9 Air Force 3010 19A Delta IV Program. 009 Transportation Security.... 171st Air Refueling Wing PAANG... 300 Tanker Rd....... Moon Township....... PA 15108 Air Force 3010 19A...

Flowfield survey over a 75 deg swept delta wing at an angle of attack of 20.5 deg

NASA Technical Reports Server (NTRS)

Kjelgaard, S. O.; Sellers, W. L., III; Weston, R. P.

1986-01-01

An experimental investigation of the flowfield over a 75 deg swept delta wing at an angle of attack of 20.5 deg has been conducted. The data include pitot pressure surveys and two types of flow visualization. Surface and flowfield visualization data were obtained at Reynolds number, Rn, ranging from 0.5 to 2.0 million in increments of 0.25 million. Detailed pitot pressure surveys were made at five longitudinal stations at Rn = 0.5, 1.0, and 1.5 million in both the primary and secondary vortices. The results indicate that Reynolds number has only a minor effect on the global structure of the flowfield in the Reynolds number range that was investigated. The boundary layer transitions from laminar to turbulent at the trailing edge of the wing at Rn = 1.0 x 10 to the 6th, and the transition moves forward to x/L = 0.4 at Rn = 2.0 x 10 to the 6th. The positions of the primary vortex cores are insensitive to Reynolds number in this range; however, the lateral position of the secondary vortex core moves outboard aft of the region where the boundary layer transitions from laminar to turbulent.

Fuzzy logic-based flight control system design

NASA Astrophysics Data System (ADS)

Nho, Kyungmoon

The application of fuzzy logic to aircraft motion control is studied in this dissertation. The self-tuning fuzzy techniques are developed by changing input scaling factors to obtain a robust fuzzy controller over a wide range of operating conditions and nonlinearities for a nonlinear aircraft model. It is demonstrated that the properly adjusted input scaling factors can meet the required performance and robustness in a fuzzy controller. For a simple demonstration of the easy design and control capability of a fuzzy controller, a proportional-derivative (PD) fuzzy control system is compared to the conventional controller for a simple dynamical system. This thesis also describes the design principles and stability analysis of fuzzy control systems by considering the key features of a fuzzy control system including the fuzzification, rule-base and defuzzification. The wing-rock motion of slender delta wings, a linear aircraft model and the six degree of freedom nonlinear aircraft dynamics are considered to illustrate several self-tuning methods employing change in input scaling factors. Finally, this dissertation is concluded with numerical simulation of glide-slope capture in windshear demonstrating the robustness of the fuzzy logic based flight control system.

Shuttle ascent and shock impingement aerodynamic heating studies

NASA Technical Reports Server (NTRS)

Lanning, W. D.; Hung, F. T.

1971-01-01

The collection and analysis of aerodynamic heating data obtained from shock impingement experimental investigation were completed. The data were categorized into four interference areas; fin leading edge, wing/fuselage fin/plate corners, and space shuttle configurations. The effects of shock impingement were found to increase the heating rates 10 to 40 times the undisturbed values. A test program was completed at NASA/Langley Research Center to investigate the magnitudes and surface patterns of the mated shock interference flowfield. A 0.0065 scale thin-skin model of the MDAC 256-20 space shuttle booster mated with a Stycast model of the MDAC Internal tank orbiter was tested in the 20-inch M=6 tunnel, the 31-inch M=10 tunnel, and the 48-inch Unitary Plan Tunnel. The gap region of the ascent configuration was the principal area of interest where both thermocouple and phase-change paint data were obtained. Pressure and heat transfer distributions data on the leeward surface of a 75-degree sweep slab delta wing are presented. The effects of surface roughness on boundary layer transition and aerodynamic heating were investigated.

Wind Tunnel Investigation of the Effects of Surface Porosity and Vertical Tail Placement on Slender Wing Vortex Flow Aerodynamics at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2007-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity and vertical tail placement on vortex flow development and interactions about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used that featured pressure sensitive paint (PSP), laser vapor screen (LVS), and schlieren, These techniques were combined with conventional electronically-scanned pressure (ESP) and six-component force and moment measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading edge extension (LEX) and the placement of centerline and twin vertical tails on the vortex-dominated flow field of a 65 cropped delta wing model. Test results were obtained at free-stream Mach numbers of 1.6, 1.8, and 2.1 and a Reynolds number per foot of 2.0 million. LEX porosity promoted a wing vortex-dominated flow field as a result of a diffusion and weakening of the LEX vortex. The redistribution of the vortex-induced suction pressures contributed to large nose-down pitching moment increments but did not significantly affect the vortex-induced lift. The trends associated with LEX porosity were unaffected by vertical tail placement. The centerline tail configuration generally provided more stable rolling moments and yawing moments compared to the twin wing-mounted vertical tails. The strength of a complex system of shock waves between the twin tails was reduced by LEX porosity.

F-16XL ship #1 crew

NASA Technical Reports Server (NTRS)

1995-01-01

November 27, 1995 Photograph of the F-16XL Ship #1 Cranked-Arrow Wing Aerodynamic Project (CAWAP) Test Team; from left to right, Ron Wilcox; Operations Engineer, Art Cope; Aircraft Mechanic, Dave Fisher; Chief Project Engineer, Dick Denman; Aircraft Mechanic, Bob Garcia; A/C Crew Chief, Susan Ligon; Aircraft Mechanic, Rodger Tarango; Mobile Operations Facility (MOF) Staff, Jerry Cousins; Aircraft Mechanic, Bruce Gallmeyer; MOF Staff, and Mike Reardon; Aircraft Mechanic/Helper. The modified airplane features a delta 'cranked-arrow' wing with strips of tubing along the leading edge to the trailing edge to sense static on the wing and obtain pressure distribution data. The right wing receives data on pressure distribution and the left wing has three types of instrumentation - preston tubes to measure local skin friction, boundary layer rakes to measure boundary layer profiles (the layer where the air interacts with the surfaces of a moving aircraft), and hot films to determine boundary layer transition locations. The first flight of CAWAP occurred at NASA's Dryden Flight Research Center, Edwards, California, on November 21, 1995, and the test program ended in April 1996.

Quantifying the Effect of Pressure Sensitive Paint On Aerodynamic Data

NASA Technical Reports Server (NTRS)

Amer, T. R.; Obara, C. J.; Liu, T.

2003-01-01

A thin pressure sensitive paint (PSP) coating can slightly modify the overall shape of a wind-tunnel model and produce surface roughness or smoothness that does not exist on the unpainted model. These undesirable changes in model geometry may alter flow over the model, and affect the pressure distribution and aerodynamic forces and moments on the model. This study quantifies the effects of PSP on three models in low-speed, transonic and supersonic flow regimes. At a 95% confidence level, the PSP effects on the integrated forces are insignificant for a slender arrow-wing-fuselage model and delta wing model with two different paints at Mach 0.2, 1.8, and 2.16 relative to the total balance accuracy limit. The data displayed a repeatability of 2.5 drag counts, while the balance accuracy limit was about 5.5 drag counts. At transonic speeds, the paint has a localized effect at high angles of attack and has a resolvable effect on the normal force, which is significant relative to the balance accuracy limit. For low speeds, the PSP coating has a localized effect on the pressure tap measurements, which leads to an appreciable decrease in the pressure tap reading. Moreover, the force and moment measurements had a poor precision, which precluded the ability to measure the PSP effect for this particular test.

Transition of the Laminar Boundary Layer on a Delta Wing with 74 degree Sweep in Free Flight at Mach Numbers from 2.8 to 5.3

NASA Technical Reports Server (NTRS)

Chapman, Gary T.

1961-01-01

The tests were conducted at Mach numbers from 2.8 to 5.3, with model surface temperatures small compared to boundary-layer recovery temperature. The effects of Mach number, temperature ratio, unit Reynolds number, leading-edge diameter, and angle of attack were investigated in an exploratory fashion. The effect of heat-transfer condition (i.e., wall temperature to total temperature ratio) and Mach number can not be separated explicitly in free-flight tests. However, the data of the present report, as well as those of NACA TN 3473, were found to be more consistent when plotted versus temperature ratio. Decreasing temperature ratio increased the transition Reynolds number. The effect of unit Reynolds number was small as was the effect of leading-edge diameter within the range tested. At small values of angle of attack, transition moved forward on the windward surface and rearward on the leeward surface. This trend was reversed at high angles of attack (6 deg to 18 deg). Possible reasons for this are the reduction of crossflow on the windward side and the influence of the lifting vortices on the leeward surface. When the transition results on the 740 delta wing were compared to data at similar test conditions for an unswept leading edge, the results bore out the results of earlier research at nearly zero heat transfer; namely, sweep causes a large reduction in the transition Reynolds number.

Numerical computation of viscous flow around bodies and wings moving at supersonic speeds

NASA Technical Reports Server (NTRS)

Tannehill, J. C.

1984-01-01

Research in aerodynamics is discussed. The development of equilibrium air curve fits; computation of hypersonic rarefield leading edge flows; computation of 2-D and 3-D blunt body laminar flows with an impinging shock; development of a two-dimensional or axisymmetric real gas blunt body code; a study of an over-relaxation procedure forthe MacCormack finite-difference scheme; computation of 2-D blunt body turbulent flows with an impinging shock; computation of supersonic viscous flow over delta wings at high angles of attack; and computation of the Space Shuttle Orbiter flowfield are discussed.

Eclipse project QF-106 and C-141A climbs out under tow on first tethered flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

TOW LAUNCH DEMONSTRATION - The Kelly Space & Technology (KST)/USAF/NASA Eclipse project's modified QF-106 climbs out under tow by a USAF C-141A on the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, hosted the project, providing engineering and facility support as well as the project pilot. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Eclipse project QF-106 and C-141A takeoff on first tethered flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

TOW ROPE TAKEOFF - The Kelly Space & Technology (KST)/USAF Eclipse project's modified QF-106 and a USAF C-141A takeoff for the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, hosted the project, providing engineering and facility support as well as the project pilot. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Eclipse project closeup of QF-106 under tow on takeoff on first flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

OFF THE GROUND - The Kelly Space & Technology (KST)/USAF/NASA Eclipse project's modified QF-106 lifts off under tow on the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, hosted the project, providing engineering and facility support as well as the project pilot. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Eclipse project closeup of QF-106 under tow on first tethered flight December 20, 1997

NASA Technical Reports Server (NTRS)

1997-01-01

The Kelly Space and Technology (KST)/USAF/NASA Eclipse project's modified QF-106 is shown under tow on the project's first tethered flight on December 20, 1997. The successful 18-minute-long flight reached an altitude of 10,000 feet. NASA's Dryden Flight Research Center, Edwards, California, is hosting the project, providing engineering and facility support as well as the project pilot, Mark Stucky. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wing loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Aerodynamic Analysis of Variable Geometry Raked Wingtips for Mid-Range Transonic Transport Aircraft

NASA Astrophysics Data System (ADS)

Jingeleski, David J.

Previous applications have shown that a wingtip treatment on a commercial airliner will reduce drag and increase fuel efficiency and the most common types of treatment are blended winglets and raked wingtips. With Boeing currently investigating novel designs for its next generation of airliners, a variable geometry raked wingtip novel control effector (VGRWT/NCE) was studied to determine the aerodynamic performance benefits over an untreated wingtip. The Boeing SUGAR design employing a truss-braced wing was selected as the baseline. Vortex lattice method (VLM) and computational fluid dynamics (CFD) software was implemented to analyze the aerodynamic performance of such a configuration applied to a next-generation, transonic, mid-range transport aircraft. Several models were created to simulate various sweep positions for the VGRWT/NCE tip, as well as a baseline model with an untreated wingtip. The majority of investigation was conducted using the VLM software, with CFD used largely as a validation of the VLM analysis. The VGRWT/NCE tip was shown to increase the lift of the wing while also decreasing the drag. As expected, the unswept VGRWT/NCE tip increases the amount of lift available over the untreated wingtip, which will be very beneficial for take-off and landing. Similarly, the swept VGRWT/NCE tip reduced the drag of the wing during cruise compared to the unmodified tip, which will favorably impact the fuel efficiency of the aircraft. Also, the swept VGRWT/NCE tip showed an increase in moment compared to the unmodified wingtip, implying an increase in stability, as well providing an avenue for roll control and gust alleviation for flexible wings. CFD analysis validated VLM as a useful low fidelity tool that yielded quite accurate results. The main results of this study are tabulated "deltas" in the forces and moments on the VGRWT/NCE tip as a function of sweep angle and aileron deflection compared to the baseline wing. A side study of the effects of the joint between the main wing and the movable tip showed that the drag impact can be kept small by careful design.

High Speed Vortex Flows

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Wilcox, Floyd J., Jr.; Bauer, Steven X. S.; Allen, Jerry M.

2000-01-01

A review of the research conducted at the National Aeronautics and Space Administration (NASA), Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data reviewed is for flat plates, cavities, bodies, missiles, wings, and aircraft. These data are presented and discussed relative to the design of future vehicles. Also presented is a brief historical review of the extensive body of high-speed vortex flow research from the 1940s to the present in order to provide perspective of the NASA LaRC's high-speed research results. Data are presented which show the types of vortex structures which occur at supersonic speeds and the impact of these flow structures to vehicle performance and control is discussed. The data presented shows the presence of both small- and large scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices and the downstream fins. It was shown that these vortex flow interference effects could be both positive and negative. Data are shown which highlights the effect that leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber has on the aerodynamics of and flow over delta wings. The observed flow fields for delta wings (i.e. separation bubble, classical vortex, vortex with shock, etc.) are discussed in the context of' aircraft design. And data have been shown that indicate that aerodynamic performance improvements are available by considering vortex flows as a primary design feature. Finally a discussing of a design approach for wings which utilize vortex flows for improved aerodynamic performance at supersonic speed is presented.

Flow past a Flat Plate with a Vortex/sink Combination

NASA Technical Reports Server (NTRS)

Mourtos, N. J.

1984-01-01

An attempt was made to model the so called leading edge vortex which forms over the leading edge of delta wings at high angles of attack. A simplified model was considered, namely that of a two-dimensional, inviscid, incompressible steady flow around a flat plate at an angle of attack with a stationary vortex detached on top, as well as a sink to simulate the strong spanwise flow. The results appear to agree qualitatively with experiments. A comparison was also made between the lift and the drag of this model and the corresponding results for two classical solutions: (1) that of totally attached flow over the plate with the Kutta condition satisfied at the trailing edge only: and (2) the Helmholtz solution of totally separated flow over the plate.

Flow past a flat plat with a vortex/sink combination

NASA Technical Reports Server (NTRS)

Mourtos, N. J.

1985-01-01

An attempt was made to model the so called leading edge vortex which forms over the leading edge of delta wings at high angles of attack. A simplified model was considered, namely that of a two-dimensional, inviscid, incompressible steady flow around a flat plate at an angle of attack with a stationary vortex detached on top, as well as a sink to simulate the strong spanwise flow. The results appear to agree qualitatively with experiments. A comparison was also made between the lift and the drag of this model and the corresponding results for two classical solutions: (1) that of totally attached flow over the plate with the Kutta condition satisfied at the trailing edge only; and (2) the Helmholtz solution of totally separated flow over the plate.

Vortex Breakdown over Slender Delta Wings (Eclatement tourbillonnaire sur les ailes delta effil es)

DTIC Science & Technology

2009-11-01

flow patterns for a) experiments of Mitchell et. al ., b) grid G9A4 fully 15-12 turbulent, c) grid G9A4 laminar to turbulent transition at 30% root...tourbillonnaires et en particulier les tourbillons de bord d’attaque subissent une désorganisation soudaine connue sous le nom de rupture du vortex. Ce...attack in the range of –10° to 36°, an amplitude of 5° to 26° and an oscillation frequency of 0.2 to 1.5 Hz. 8) TPI Test Case De Luca et al . tested a

Experimental study of delta wing leading-edge devices for drag reduction at high lift

NASA Technical Reports Server (NTRS)

Johnson, T. D., Jr.; Rao, D. M.

1982-01-01

The drag reduction devices selected for evaluation were the fence, slot, pylon-type vortex generator, and sharp leading-edge extension. These devices were tested on a 60 degree flatplate delta (with blunt leading edges) in the Langley Research Center 7- by 10-foot high-speed tunnel at low speed and to angles of attack of 28 degrees. Balance and static pressure measurements were taken. The results indicate that all the devices had significant drag reduction capability and improved longitudinal stability while a slight loss of lift and increased cruise drag occurred.

Investigation of the Vortex Tab. M.S. Thesis

NASA Technical Reports Server (NTRS)

Hoffler, K. D.

1985-01-01

An investigation was made into the drag reduction capability of vortex tabs on delta wing vortex flaps. The vortex tab is an up-deflected leading edge portion of the vortex flap. Tab deflection augments vortex suction on the flap, thus improving its thrust, but the tab itself is drag producing. Whether a net improvement in the drag reduction can be obtained with vortex tabs, in comparison with plane vortex flaps of the same total area, was the objective of this investigation. Wind tunnel tests were conducted on two models, and analytical studies were performed on one of them using a free vortex sheet theory.

Impact of flow unsteadiness on maneuvers and loads of agile aircraft

NASA Technical Reports Server (NTRS)

Jarrah, M. Ameen; Ashley, Holt

1989-01-01

A program of airload measurements on a family of low-aspect-ratio delta wings with sharp leading edges, subjected to large amplitude pitch transients with angles of attack up to 90 deg, is reviewed. Even for small values of the pitch-rate parameter, representative of maneuvers anticipated for agile aircraft, the force and moment overshoots can exceed by 50 percent their steady-state values. This is explained in terms of the hysteretic behavior of the breakdown locations of leading-edge vortices. An approximate theoretical model is proposed which includes the breakdown hysteresis as part of a three-term representation of the unsteady chordwise load distribution.

Identification of aerodynamic models for maneuvering aircraft

NASA Technical Reports Server (NTRS)

Lan, C. Edward; Hu, C. C.

1992-01-01

A Fourier analysis method was developed to analyze harmonic forced-oscillation data at high angles of attack as functions of the angle of attack and its time rate of change. The resulting aerodynamic responses at different frequencies are used to build up the aerodynamic models involving time integrals of the indicial type. An efficient numerical method was also developed to evaluate these time integrals for arbitrary motions based on a concept of equivalent harmonic motion. The method was verified by first using results from two-dimensional and three-dimensional linear theories. The developed models for C sub L, C sub D, and C sub M based on high-alpha data for a 70 deg delta wing in harmonic motions showed accurate results in reproducing hysteresis. The aerodynamic models are further verified by comparing with test data using ramp-type motions.

Analysis of Asymmetric Aircraft Aerodynamics Due to an Experimental Wing Glove

NASA Technical Reports Server (NTRS)

Hartshorn, Fletcher

2011-01-01

Aerodynamic computational fluid dynamics analysis of a wing glove attached to one wing of a business jet is presented and discussed. A wing glove placed on only one wing will produce asymmetric aerodynamic effects that will result in overall changes in the forces and moments acting on the aircraft. These changes, referred to as deltas, need to be determined and quantified to ensure that the wing glove does not have a significant effect on the aircraft flight characteristics. TRANAIR (Calmar Research Corporation, Cato, New York), a nonlinear full potential solver, and Star-CCM+ (CD-adapco, Melville, New York), a finite volume full Reynolds-averaged Navier-Stokes computational fluid dynamics solver, are used to analyze a full aircraft with and without the glove at a variety of flight conditions, aircraft configurations, and angles of attack and sideslip. Changes in the aircraft lift, drag, and side force along with roll, pitch, and yaw are presented. Span lift and moment distributions are also presented for a more detailed look at the effects of the glove on the aircraft. Aerodynamic flow phenomena due to the addition of the glove are discussed. Results show that the glove produces only small changes in the aerodynamic forces and moments acting on the aircraft, most of which are insignificant.

Experimental and numerical research on the aerodynamics of unsteady moving aircraft

NASA Astrophysics Data System (ADS)

Bergmann, Andreas; Huebner, Andreas; Loeser, Thomas

2008-02-01

For the experimental determination of the dynamic wind tunnel data, a new combined motion test capability was developed at the German-Dutch Wind Tunnels DNW for their 3 m Low Speed Wind Tunnel NWB in Braunschweig, Germany, using a unique six degree-of-freedom test rig called ‘Model Positioning Mechanism’ (MPM) as an improved successor to the older systems. With that cutting-edge device, several transport aircraft configurations including a blended wing body configuration were tested in different modes of oscillatory motions roll, pitch and yaw as well as delta-wing geometries like X-31 equipped with remote controlled rudders and flaps to be able to simulate realistic flight maneuvers, e.g., a Dutch Roll. This paper describes the motivation behind these tests and the test setup and in addition gives a short introduction into time accurate maneuver-testing capabilities incorporating models with remote controlled control surfaces. Furthermore, the adaptation of numerical methods for the prediction of dynamic derivatives is described and some examples with the DLR-F12 configuration will be given. The calculations are based on RANS-solution using the finite volume parallel solution algorithm with an unstructured discretization concept (DLR TAU-code).

Aerodynamic control of NASP-type vehicles through vortex manipulation, volume 4

NASA Technical Reports Server (NTRS)

Smith, Brooke C.; Suarez, Carlos J.; Porada, William M.; Malcolm, Gerald N.

1993-01-01

Forebody Vortex Control (FVC) is an emerging technology that has received widespread and concentrated attention by many researchers for application on fighter aircraft to enhance aerodynamic controllability at high angles of attack. This research explores potential application of FVC to a NASP-type configuration. The configuration investigated is characterized by a slender, circular cross-section forebody and a 78 deg swept delta wing. A man-in-the-loop, six-degress-of-freedom, high-fidelity simulation was developed that demonstrates the implementation and advantages of pneumatic forebody vortex control. Static wind tunnel tests were used as the basis for the aerodynamic characteristics modeled in the simulation. Dynamic free-to-roll wind tunnel tests were analyzed and the wing rock motion investigated. A non-linear model of the dynamic effects of the bare airframe and the forebody vortex control system were developed that closely represented the observed behavior. Multiple state-of-the-art digital flight control systems were developed that included different utilizations of pneumatic vortex control. These were evaluated through manned simulation. Design parameters for a pneumatic forebody vortex control system were based on data collected regarding the use of blowing and the mass flow required during realistic flight maneuvers.

Effect of Tail Dihedral on Lateral Control Effectiveness at High Subsonic Speeds of Differentially Deflected Horizontal-Tail Surfaces on a Configuration having a Thin Highly Tapered Wing

NASA Technical Reports Server (NTRS)

Fournier, Paul G.

1959-01-01

Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel to determine the effect of tail dihedral on lateral control effectiveness of a complete-model configuration having differentially deflected horizontal-tail surfaces. Limited tests were made to determine the lateral characteristics as well as the longitudinal characteristics in sideslip. The wing had an aspect ratio of 3, a taper ratio of 0.14, 28.80 deg sweep of the quarter-chord line with zero sweep at the 80-percent-chord line, and NACA 65A004 airfoil sections. The test Mach number range extended from 0.60 to 0.92. There are only small variations in the roll effectiveness parameter C(sub iota delta) with negative tail dihedral angle. The tail size used on the test model, however, is perhaps inadequate for providing the roll rates specified by current military requirements at subsonic speeds. The lateral aerodynamic characteristics were essentially constant throughout the range of sideslip angle from 12 deg to -12 deg. A general increase in yawing moment was noted with increased negative dihedral throughout the Mach number range.

Adaptive Suction and Blowing for Twin-Tail Buffet Control

NASA Technical Reports Server (NTRS)

Kandil, Osama A.; Yang, Zhi

1999-01-01

Adaptive active flow control for twin-tail buffet alleviation is investigated. The concept behind this technique is to place control ports on the tail outer and inner surfaces with flow suction or blowing applied through these ports in order to minimize the pressure difference across the tail. The suction or blowing volume flow rate from each port is proportional to the pressure difference across the tail at this location. A parametric study of the effects of the number and location of these ports on the buffet response is carried out. The computational model consists of a sharp-edged delta wing of aspect ratio one and swept-back flexible twin tail with taper ratio of 0.23. This complex multidisciplinary problem is solved sequentially using three sets of equations for the fluid flow, aeroelastic response and grid deformation, using a dynamic multi-block grid structure. The computational model is pitched at 30 deg angle of attack. The freestream Mach number and Reynolds number are 0.3 and 1.25 million, respectively. The model is investigated for the inboard position of the twin tails, which corresponds to a separation distance between the twin tails of 33% of the wing span. Comparison of the time history and power spectral density responses of the tails for various distributions of the control ports are presented and discussed.

Supersonic aerodynamic characteristics of a circular body Earth-to-Orbit vehicle

NASA Technical Reports Server (NTRS)

Ware, George M.; Engelund, Walter C.; Macconochie, Ian O.

1994-01-01

The circular body configuration is a generic single- or multi-stage reusable Earth-to-orbit transport. A thick clipped-delta wing is the major lifting surface. For directional control, three different vertical fin arrangements were investigated: a conventional aft-mounted center fin, wingtip fins, and a nose-mounted fin. The tests were conducted in the Langley Unitary Plan Wind Tunnel. The configuration is longitudinally stable about the estimated center of gravity of 0.72 body length up to a Mach number of about 3.0. Above Mach 3.0, the model is longitudinally unstable at low angles of attack but has a stable secondary trim point at angles of attack above 30 deg. The model has sufficient pitch control authority with elevator and body flap to produce stable trim over the test range. The model with the center fin is directionally stable at low angles of attack up to a Mach number of 3.90. The rudder-like surfaces on the tip fins and the all-movable nose fin are designed as active controls to produce artificial directional stability and are effective in producing yawing moment. The wing trailing-edge aileron surfaces are effective in producing rolling moment, but they also produce large adverse yawing moment.

Aerodynamic prediction techniques for hypersonic configuration design

NASA Technical Reports Server (NTRS)

1981-01-01

An investigation of approximate theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds was performed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Potential theory was examined in detail to meet this objective. Numerical pilot codes were developed for relatively simple three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the computations indicate good agreement with higher order solutions and experimental results for a variety of wing, body, and wing-body shapes for values of the hypersonic similarity parameter M delta approaching one.

Comparison of measured and computed pitot pressures in a leading edge vortex from a delta wing

NASA Technical Reports Server (NTRS)

Murman, Earll M.; Powell, Kenneth G.

1987-01-01

Calculations are presented for a 75-deg swept flat plate wing tested at a freestream Mach number of 1.95 and 10 degrees angle of attack. Good agreement is found between computational data and previous experimental pitot pressure measurements in the core of the vortex, suggesting that the total pressure losses predicted by the Euler equation solvers are not errors, but realistic predictions. Data suggest that the magnitude of the total pressure loss is related to the circumferential velocity field through the vortex, and that it increases with angle of attack and varies with Mach number and sweep angle.

Adaptive wing and flow control technology

NASA Astrophysics Data System (ADS)

Stanewsky, E.

2001-10-01

The development of the boundary layer and the interaction of the boundary layer with the outer “inviscid” flow field, exacerbated at high speed by the occurrence of shock waves, essentially determine the performance boundaries of high-speed flight. Furthermore, flight and freestream conditions may change considerably during an aircraft mission while the aircraft itself is only designed for multiple but fixed design points thus impairing overall performance. Consequently, flow and boundary layer control and adaptive wing technology may have revolutionary new benefits for take-off, landing and cruise operating conditions for many aircraft by enabling real-time effective geometry optimization relative to the flight conditions. In this paper we will consider various conventional and novel means of boundary layer and flow control applied to moderate-to-large aspect ratio wings, delta wings and bodies with the specific objectives of drag reduction, lift enhancement, separation suppression and the improvement of air-vehicle control effectiveness. In addition, adaptive wing concepts of varying complexity and corresponding aerodynamic performance gains will be discussed, also giving some examples of possible structural realizations. Furthermore, penalties associated with the implementation of control and adaptation mechanisms into actual aircraft will be addressed. Note that the present contribution is rather application oriented.

ASYMMETRIC ABSORPTION PROFILES OF Ly{alpha} AND Ly{beta} IN DAMPED Ly{alpha} SYSTEMS

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lee, Hee-Won, E-mail: hwlee@sejong.ac.kr

2013-08-01

Damped Ly{alpha} systems observed in the quasar spectra are characterized by a high neutral hydrogen column density, N{sub HI} > 2 x 10{sup 20} cm{sup -2}. The absorption wing profiles are often fitted using the Voigt function due to the fact that the scattering cross section near the resonant line center is approximately described by the Lorentzian function. Since a hydrogen atom has infinitely many p states that participate in the electric dipole interaction, the cross section starts to deviate from the Lorentzian in an asymmetric way in the line wing regions. We investigate this asymmetry in the absorption linemore » profiles around Ly{alpha} and Ly{beta} as a function of the neutral hydrogen column density N{sub HI}. In terms of {Delta}{lambda} {identical_to} {lambda} - {lambda}{sub {alpha}}, we expand the Kramers-Heisenberg formula around Ly{alpha} to find {sigma}({lambda}) {approx_equal} (0.5f{sub 12}){sup 2}{sigma}{sub T}({Delta}{lambda}/{lambda}{sub {alpha}}){sup -2}[1 + 3.792({Delta}{lambda}/{lambda}{sub {alpha}})], where f{sub 12} and {sigma}{sub T} are the oscillator strength of Ly{alpha} and the Thomson scattering cross section, respectively. In terms of {Delta}{lambda}{sub 2} {identical_to} {lambda} - {lambda}{sub {beta}} in the vicinity of Ly{beta}, the total scattering cross section, given as the sum of cross sections for Rayleigh and Raman scattering, is shown to be {sigma}({lambda}) {approx_equal} {sigma}{sub T}(0.5f{sub 13}){sup 2}(1 + R{sub 0})({Delta}{lambda}{sub 2}/{lambda}{sub {beta}}){sup -2}[1 - 24.68({Delta}{lambda}{sub 2}/{lambda}{sub {beta}})] with f{sub 13} and the factor R{sub 0} = 0.1342 being the oscillator strength for Ly{beta} and the ratio of the Raman cross section to Rayleigh cross section, respectively. A redward asymmetry develops around Ly{alpha}, whereas a blue asymmetry is obtained for Ly{beta}. The absorption center shifts are found to be almost proportional to the neutral hydrogen column density.« less

Eclipse program C-141A aircraft

NASA Technical Reports Server (NTRS)

1997-01-01

This photograph shows the Air Force C-141A that was used in the Eclipse project as a tow vehicle. In 1997 and 1998, the Dryden Flight Research Center at Edwards, California, supported and hosted a Kelly Space & Technology, Inc. project called Eclipse, which sought to demonstrate the feasibility of a reusable tow-launch vehicle concept. The project goal was to successfully tow, inflight, a modified QF-106 delta-wing aircraft with an Air Force C-141A transport aircraft. This would demonstrate the possibility of towing and launching an actual launch vehicle from behind a tow plane. Dryden was the responsible test organization and had flight safety responsibility for the Eclipse project. Dryden provided engineering, instrumentation, simulation, modification, maintenance, range support, and research pilots for the test program. The Air Force Flight Test Center (AFFTC), Edwards, California, supplied the C-141A transport aircraft and crew and configured the aircraft as needed for the tests. The AFFTC also provided the concept and detail design and analysis as well as hardware for the tow system and QF-106 modifications. Dryden performed the modifications to convert the QF-106 drone into the piloted EXD-01 (Eclipse eXperimental Demonstrator-01) experimental aircraft. Kelly Space & Technology hoped to use the results gleaned from the tow test in developing a series of low-cost, reusable launch vehicles. These tests demonstrated the validity of towing a delta-wing aircraft having high wind loading, validated the tow simulation model, and demonstrated various operational procedures, such as ground processing of in-flight maneuvers and emergency abort scenarios.

Aerodynamic characteristics of a hypersonic research airplane concept having a 70 degree swept double delta wing at Mach numbers from 1.50 to 2.86

NASA Technical Reports Server (NTRS)

Penland, J. A.; Fournier, R. H.; Marcum, D. C., Jr.

1975-01-01

An experimental investigation of the static longitudinal, lateral, and directional stability characteristics of a hypersonic research airplane concept having a 70 deg swept double-delta wing was conducted in the Langley unitary plan wind tunnel. The configuration variables included wing planform, tip fins, center fin, and scramjet engine modules. The investigation was conducted at Mach numbers from 1.50 to 2.86 and at a constant Reynolds number, based on fuselage length, of 3,330,000. Tests were conducted through an angle-of-attack range from about -4 deg to 24 deg with angles of sideslip of 0 deg and 3 deg and at elevon deflections of 0, -10, and -20 deg. The complete configuration was trimmable up to angles of attack of about 22 deg with the exception of regions at low angles of attack where positive elevon deflections should provide trim capability. The angle-of-attack range for which static longitudinal stability also exists was reduced at the higher Mach numbers due to the tendency of the complete configuration to pitch up at the higher angles of attack. The complete configuration was statically stable directionally up to trimmed angles of attack of at least 20 deg for all Mach numbers M with the exception of a region near 4 deg at M = 2.86 and exhibited positive effective dihedral at all positive trimmed angles of attack.

Aerodynamic shape optimization directed toward a supersonic transport using sensitivity analysis

NASA Technical Reports Server (NTRS)

Baysal, Oktay

1995-01-01

This investigation was conducted from March 1994 to August 1995, primarily, to extend and implement the previously developed aerodynamic design optimization methodologies for the problems related to a supersonic transport design. These methods had demonstrated promise to improve the designs (more specifically, the shape) of aerodynamic surfaces, by coupling optimization algorithms (OA) with Computational Fluid Dynamics (CFD) algorithms via sensitivity analyses (SA) with surface definition methods from Computer Aided Design (CAD). The present extensions of this method and their supersonic implementations have produced wing section designs, delta wing designs, cranked-delta wing designs, and nacelle designs, all of which have been reported in the open literature. Despite the fact that these configurations were highly simplified to be of any practical or commercial use, they served the algorithmic and proof-of-concept objectives of the study very well. The primary cause for the configurational simplifications, other than the usual simplify-to-study the fundamentals reason, were the premature closing of the project. Only after the first of the originally intended three-year term, both the funds and the computer resources supporting the project were abruptly cut due to their severe shortages at the funding agency. Nonetheless, it was shown that the extended methodologies could be viable options in optimizing the design of not only an isolated single-component configuration, but also a multiple-component configuration in supersonic and viscous flow. This allowed designing with the mutual interference of the components being one of the constraints all along the evolution of the shapes.

Determination of space shuttle flow field by the three-dimensional method of characteristics

NASA Technical Reports Server (NTRS)

Chu, C.; Powers, S. A.

1972-01-01

The newly improved three-dimensional method of characteristics program has been applied successfully to the calculation of flow fields over a variety of bodies including slab delta wings and shuttle orbiters. Flow fields over fuselage shapes for Mach numbers as low as 1.5 have been calculated. Some typical results are presented.

A Gaussian measure of quantum phase noise

NASA Technical Reports Server (NTRS)

Schleich, Wolfgang P.; Dowling, Jonathan P.

1992-01-01

We study the width of the semiclassical phase distribution of a quantum state in its dependence on the average number of photons (m) in this state. As a measure of phase noise, we choose the width, delta phi, of the best Gaussian approximation to the dominant peak of this probability curve. For a coherent state, this width decreases with the square root of (m), whereas for a truncated phase state it decreases linearly with increasing (m). For an optimal phase state, delta phi decreases exponentially but so does the area caught underneath the peak: all the probability is stored in the broad wings of the distribution.

Numerical solutions of Navier-Stokes equations for a Butler wing

NASA Technical Reports Server (NTRS)

Abolhassani, J. S.; Tiwari, S. N.

1985-01-01

The flow field is simulated on the surface of a given delta wing (Butler wing) at zero incident in a uniform stream. The simulation is done by integrating a set of flow field equations. This set of equations governs the unsteady, viscous, compressible, heat conducting flow of an ideal gas. The equations are written in curvilinear coordinates so that the wing surface is represented accurately. These equations are solved by the finite difference method, and results obtained for high-speed freestream conditions are compared with theoretical and experimental results. In this study, the Navier-Stokes equations are solved numerically. These equations are unsteady, compressible, viscous, and three-dimensional without neglecting any terms. The time dependency of the governing equations allows the solution to progress naturally for an arbitrary initial initial guess to an asymptotic steady state, if one exists. The equations are transformed from physical coordinates to the computational coordinates, allowing the solution of the governing equations in a rectangular parallel-piped domain. The equations are solved by the MacCormack time-split technique which is vectorized and programmed to run on the CDC VPS 32 computer.

Analytical Investigation of a Flicker-Type Roll Control for a Mach Number 6 Missile with Aerodynamic Controls Over An Altitude Range of 82,000 to 282,000 feet

NASA Technical Reports Server (NTRS)

Lundstrom, Reginald R.; Whitman, Ruth I.

1959-01-01

An analytical investigation has been carried out to determine the responses of a flicker-type roll control incorporated in a missile which traverses a range of Mach number of 6.3 at an altitude of 82,000 feet to 5.26 at an altitude of 282,000 feet. The missile has 80 deg delta wings in a cruciform arrangement with aerodynamic controls attached to the fuselage near the wing trailing edge and indexed 450 to the wings. Most of the investigation was carried out on an analog computer. Results showed that roll stabilization that may be adequate for many cases can be obtained over the altitude range considered, providing that the rate factor can be changed with altitude. The response would be improved if the control deflection were made larger at the higher altitudes. lag times less than 0.04 second improve the response appreciably. Asymmetries that produce steady rolling moments can be very detrimental to the response in some cases. The wing damping made a negligible contribution to the response.

Heat transfer tests of the McDonnell-Douglas delta wing orbiter mated with -17A booster at Mach number 8

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

A wind tunnel test program to determine aerodynamic heat transfer distributions on the McDonnell-Douglas configurations is reported. The tests were conducted at the Arnold Engineering Development Center (AEDC) in Tunnel B of the von Karman Gas Dynamics Facility (VKF). Heat-transfer rates were determined by the phase-change paint technique on 0.011-scale Stycast models using Tempilaq as the surface temperature indicator. The nominal test conditions were; Mach 8, freestream unit Reynolds numbers of 0.8 x one million, 2.5 x one million, and 3.7 x one million, and angles of attack of -5 deg, 0 deg, +5deg. Model details, test conditions, phase-change paint photographs and reduced heat-transfer coefficients are presented.

Rotor Design Options for Improving XV-15 Whirl-Flutter Stability Margins

NASA Technical Reports Server (NTRS)

Acree, C. W., Jr.; Peyran, R. J.; Johnson, Wayne

2004-01-01

Rotor design changes intended to improve tiltrotor whirl-flutter stability margins were analyzed. A baseline analytical model of the XV-15 was established, and then a thinner, composite wing was designed to be representative of a high-speed tiltrotor. The rotor blade design was modified to increase the stability speed margin for the thin-wing design. Small rearward offsets of the aerodynamic-center locus with respect to the blade elastic axis created large increases in the stability boundary. The effect was strongest for offsets at the outboard part of the blade, where an offset of the aerodynamic center by 10% of tip chord improved the stability margin by over 100 knots. Forward offsets of the blade center of gravity had similar but less pronounced effects. Equivalent results were seen for swept-tip blades. Appropriate combinations of sweep and pitch stiffness completely eliminated whirl flutter within the speed range examined; alternatively, they allowed large increases in pitch-flap coupling (delta-three) for a given stability margin. A limited investigation of the rotor loads in helicopter and airplane configuration showed only minor increases in loads.

Control of Interacting Vortex Flows at Subsonic and Transonic Speeds Using Passive Porosity

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2003-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) 8-foot Transonic Pressure Tunnel (TPT) to determine the effects of passive surface porosity on vortex flow interactions about a general research fighter configuration at subsonic and transonic speeds. Flow- through porosity was applied to a wind leading-edge extension (LEX) mounted to a 65 deg cropped delta wind model to promote large nose-down pitching moment increments at high angles of attack. Porosity decreased the vorticity shed from the LEX, which weakened the LEX vortex and altered the global interactions of the LEX and wing vortices at high angles of attack. Six-component forces and moments and wing upper surface static pressure distributions were obtained at free- stream Mach numbers of 0.50, 0.85, and 1.20, Reynolds number of 2.5(10(exp-6) per foot, angles of attack up to 30 deg and angles of sideslip to plus or minus 8 deg. The off-surface flow field was visualized in selected cross-planes using a laser vapor screen flow visualization technique. Test data were obtained with a centerline vertical tail and with alternate twin, wing-mounted vertical fins having 0 deg and 30 deg cant angles. In addition, the porosity of the LEX was compartmentalized to determine the sensitivity of the vortex- dominated aerodynamics to the location and level of porosity applied to the LEX.

Analysis of a two-dimensional type 6 shock-interference pattern using a perfect-gas code and a real-gas code

NASA Technical Reports Server (NTRS)

Bertin, J. J.; Graumann, B. W.

1973-01-01

Numerical codes were developed to calculate the two dimensional flow field which results when supersonic flow encounters double wedge configurations whose angles are such that a type 4 pattern occurs. The flow field model included the shock interaction phenomena for a delta wing orbiter. Two numerical codes were developed, one which used the perfect gas relations and a second which incorporated a Mollier table to define equilibrium air properties. The two codes were used to generate theoretical surface pressure and heat transfer distributions for velocities from 3,821 feet per second to an entry condition of 25,000 feet per second.

Wind Tunnel Investigation of Passive Vortex Control and Vortex-Tail Interactions on a Slender Wing at Subsonic and Transonic Speeds

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2013-01-01

A wind tunnel experiment was conducted in the NASA Langley 8-Foot Transonic Pressure Tunnel to determine the effects of passive porosity on vortex flow interactions about a slender wing configuration at subsonic and transonic speeds. Flow-through porosity was applied in several arrangements to a leading-edge extension, or LEX, mounted to a 65-degree cropped delta wing as a longitudinal instability mitigation technique. Test data were obtained with LEX on and off in the presence of a centerline vertical tail and twin, wing-mounted vertical fins to quantify the sensitivity of the aerodynamics to tail placement and orientation. A close-coupled canard was tested as an alternative to the LEX as a passive flow control device. Wing upper surface static pressure distributions and six-component forces and moments were obtained at Mach numbers of 0.50, 0.85, and 1.20, unit Reynolds number of 2.5 million, angles of attack up to approximately 30 degrees, and angles of sideslip to +/-8 degrees. The off-surface flow field was visualized in cross planes on selected configurations using a laser vapor screen flow visualization technique. Tunnel-to-tunnel data comparisons and a Reynolds number sensitivity assessment were also performed. 15.

Application of fully stressed design procedures to redundant and non-isotropic structures

NASA Technical Reports Server (NTRS)

Adelman, H. M.; Haftka, R. T.; Tsach, U.

1980-01-01

An evaluation is presented of fully stressed design procedures for sizing highly redundant structures including structures made of composite materials. The evaluation is carried out by sizing three structures: a simple box beam of either composite or metal construction; a low aspect ratio titanium wing; and a titanium arrow wing for a conceptual supersonic cruise aircraft. All three structures are sized by ordinary fully-stressed design (FSD) and thermal fully stressed design (TFSD) for combined mechanical and thermal loads. Where possible, designs are checked by applying rigorous mathematical programming techniques to the structures. It is found that FSD and TFSD produce optimum designs for the metal box beam, but produce highly non-optimum designs for the composite box beam. Results from the delta wing and arrow wing indicate that FSD and TFSD exhibits slow convergence for highly redundant metal structures. Further, TFSD exhibits slow oscillatory convergence behavior for the arrow wing for very high temperatures. In all cases where FSD and TFSD perform poorly either in obtaining nonoptimum designs or in converging slowly, the assumptions on which the algorithms are based are grossly violated. The use of scaling, however, is found to be very effective in obtaining fast convergence and efficiently produces safe designs even for those cases when FSD and TFSD alone are ineffective.

The EGF Repeat-Specific O-GlcNAc-Transferase Eogt Interacts with Notch Signaling and Pyrimidine Metabolism Pathways in Drosophila

PubMed Central

Müller, Reto; Jenny, Andreas; Stanley, Pamela

2013-01-01

The O-GlcNAc transferase Eogt modifies EGF repeats in proteins that transit the secretory pathway, including Dumpy and Notch. In this paper, we show that the Notch ligands Delta and Serrate are also substrates of Eogt, that mutation of a putative UDP-GlcNAc binding DXD motif greatly reduces enzyme activity, and that Eogt and the cytoplasmic O-GlcNAc transferase Ogt have distinct substrates in Drosophila larvae. Loss of Eogt is larval lethal and disrupts Dumpy functions, but does not obviously perturb Notch signaling. To identify novel genetic interactions with eogt, we investigated dominant modification of wing blister formation caused by knock-down of eogt. Unexpectedly, heterozygosity for several members of the canonical Notch signaling pathway suppressed wing blister formation. And importantly, extensive genetic interactions with mutants in pyrimidine metabolism were identified. Removal of pyrimidine synthesis alleles suppressed wing blister formation, while removal of uracil catabolism alleles was synthetic lethal with eogt knock-down. Therefore, Eogt may regulate protein functions by O-GlcNAc modification of their EGF repeats, and cellular metabolism by affecting pyrimidine synthesis and catabolism. We propose that eogt knock-down in the wing leads to metabolic and signaling perturbations that increase cytosolic uracil levels, thereby causing wing blister formation. PMID:23671640

Early Program Development

NASA Image and Video Library

1961-01-01

A Dyna-Soar (Dynamic Soaring) vehicle clears the launch tower atop an Air Force Titan II launch vehicle in this 1961 artist's concept. Originally conceived by the U.S. Air Force in 1957 as a marned, rocket-propelled glider in a delta-winged configuration, the Dyna-Soar was considered by Marshall Space Flight Center planners as an upper stage for the Saturn C-2 launch vehicle.

Aerodynamic Characteristics of a Revised Target Drone Vehicle at Mach Numbers from 1.60 to 2.86

NASA Technical Reports Server (NTRS)

Blair, A. B., Jr.; Babb, C. Donald

1968-01-01

An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a revised target drone vehicle through a Mach number range from 1.60 to 2.86. The vehicle had canard surfaces and a swept clipped-delta wing with twin tip-mounted vertical tails.

Force production and time-averaged flow structure around thin, non-slender delta wings

NASA Astrophysics Data System (ADS)

Tu, Han; Green, Melissa

2017-11-01

Experimental force measurement and time-averaged three dimensional flow visualization of low Reynolds number baseline cases have been carried out on a steady flat plate delta wing. Current data will serve as steady reference for future unsteady flow and actuation cases. The comprehensive study will compare force production in highly unsteady environments, which is necessary to consider in unmanned combat aerial vehicle (UCAV) control strategies. Force measurements are carried out at angles of attack 10, 15, 20, 25 and 30 degrees. The coefficient of drag increases with angle of attack, while the coefficient of lift reaches a maximum value at 20 degrees. Time-averaged flow visualization conducted at angles of attack of 20, 25 and 30 degrees shows vortices with larger magnitude that persist farther into wake are generated at higher angles of attack. These results compare analogously with similar steady baseline experiment results of high Reynolds number conducted by collaborators. This work was supported by the Office of Naval Research under ONR Award No. N00014-16-1-2732. We also acknowledge the collaborative support of Dr. David Rival and Mr. Matthew Marzanek at Queen's University.

Space shuttle: Longitudinal and lateral aerodynamic characteristics of the 0.0035-scale GD/C aerospace booster (B-15B-1)

NASA Technical Reports Server (NTRS)

Debevoise, J. M.; Mcginnis, R. F.

1972-01-01

Force tests on a 0.0035-scale model of the General Dynamics/Convair space shuttle B-15B-1 booster were conducted in the MSFC trisonic wind tunnel during February and March 1971. Longitudinal and lateral characteristics were obtained at Mach numbers from 0.6 to 4.96. The configuration tested had a low delta wing, all-movable canard controls of delta planform, and a single vertical tail. Most of the test was devoted to obtaining data relevant to the transition from atmospheric reentry to subsonic cruise. In that portion of the test the angles of attack ranged from 6 degrees to 60 degrees, and yaw runs were made at angles of attack of 15 and 35 degrees. The rest of the test was devoted to obtaining booster-alone buildup data relevant to the launch phase. For the launch phase, the Mach number range was from 0.6 to 2.0, the angles of attack were from -10 to +10 degrees, and yaw runs were made at zero angle of attack.

KSC00padig072

NASA Image and Video Library

2000-11-02

Center Director Roy Bridges addresses the audience at the commissioning of a new high-pressure helium pipeline at Kennedy Space Center that will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS); Col. Samuel Dick, representative of the 45th Space Wing; Ramon Lugo, acting executive director, JPMO; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS

KSC-00padig072

NASA Image and Video Library

2000-11-02

Center Director Roy Bridges addresses the audience at the commissioning of a new high-pressure helium pipeline at Kennedy Space Center that will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS); Col. Samuel Dick, representative of the 45th Space Wing; Ramon Lugo, acting executive director, JPMO; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS

KSC-00padig073

NASA Image and Video Library

2000-11-02

Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS) presents an award of appreciation to H.T. Everett, KSC Propellants manager, at the commissioning of a new high-pressure helium pipeline at Kennedy Space Center. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Center Director Roy Bridges;); Col. Samuel Dick, representative of the 45th Space Wing; Ramon Lugo, acting executive director, JPMO; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS

KSC-00padig071

NASA Image and Video Library

2000-11-02

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, Ramon Lugo, acting executive director, JPMO , presents a plaque to Center Director Roy Bridges. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS); Col. Samuel Dick, representative of the 45th Space Wing; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad

KSC-00padig070

NASA Image and Video Library

2000-11-02

Jerry Jorgensen welcomes the audience to the commissioning of a new high-pressure helium pipeline at Kennedy Space Center. Jorgensen, with Space Gateway Support (SGS), is the pipeline project manager. To the right is Ramon Lugo, acting executive director, JPMO. Others at the ceremony were Center Director Roy Bridges; Col. Samuel Dick, representative of the 45th Space Wing; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad

KSC00padig071

NASA Image and Video Library

2000-11-02

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, Ramon Lugo, acting executive director, JPMO , presents a plaque to Center Director Roy Bridges. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS); Col. Samuel Dick, representative of the 45th Space Wing; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad

KSC00padig073

NASA Image and Video Library

2000-11-02

Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS) presents an award of appreciation to H.T. Everett, KSC Propellants manager, at the commissioning of a new high-pressure helium pipeline at Kennedy Space Center. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Center Director Roy Bridges;); Col. Samuel Dick, representative of the 45th Space Wing; Ramon Lugo, acting executive director, JPMO; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS

Commissioning of a new helium pipeline

NASA Technical Reports Server (NTRS)

2000-01-01

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, Ramon Lugo, acting executive director, JPMO , presents a plaque to Center Director Roy Bridges. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS); Col. Samuel Dick, representative of the 45th Space Wing; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch's worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad.

Commissioning of a new helium pipeline

NASA Technical Reports Server (NTRS)

2000-01-01

Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS) presents an award of appreciation to H.T. Everett, KSC Propellants manager, at the commissioning of a new high-pressure helium pipeline at Kennedy Space Center. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch's worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Center Director Roy Bridges;); Col. Samuel Dick, representative of the 45th Space Wing; Ramon Lugo, acting executive director, JPMO; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS.

Commissioning of a new helium pipeline

NASA Technical Reports Server (NTRS)

2000-01-01

Jerry Jorgensen welcomes the audience to the commissioning of a new high-pressure helium pipeline at Kennedy Space Center. Jorgensen, with Space Gateway Support (SGS), is the pipeline project manager. To the right is Ramon Lugo, acting executive director, JPMO. Others at the ceremony were Center Director Roy Bridges; Col. Samuel Dick, representative of the 45th Space Wing; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS. The pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile-long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch's worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad.

Commissioning of a new helium pipeline

NASA Technical Reports Server (NTRS)

2000-01-01

Center Director Roy Bridges addresses the audience at the commissioning of a new high-pressure helium pipeline at Kennedy Space Center that will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. The nine-mile- long buried pipeline will also serve as a backup helium resource for Shuttle launches. Nearly one launch's worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS); Col. Samuel Dick, representative of the 45th Space Wing; Ramon Lugo, acting executive director, JPMO; David Herst, director, Delta IV Launch Sites; Pierre Dufour, president and CEO, Air Liquide America Corporation; and Michael Butchko, president, SGS.

Aerodynamic comparison of a butterfly-like flapping wing-body model and a revolving-wing model

NASA Astrophysics Data System (ADS)

Suzuki, Kosuke; Yoshino, Masato

2017-06-01

The aerodynamic performance of flapping- and revolving-wing models is investigated by numerical simulations based on an immersed boundary-lattice Boltzmann method. As wing models, we use (i) a butterfly-like model with a body and flapping-rectangular wings and (ii) a revolving-wing model with the same wings as the flapping case. Firstly, we calculate aerodynamic performance factors such as the lift force, the power, and the power loading of the two models for Reynolds numbers in the range of 50-1000. For the flapping-wing model, the power loading is maximal for the maximum angle of attack of 90°, a flapping amplitude of roughly 45°, and a phase shift between the flapping angle and the angle of attack of roughly 90°. For the revolving-wing model, the power loading peaks for an angle of attack of roughly 45°. In addition, we examine the ground effect on the aerodynamic performance of the revolving-wing model. Secondly, we compare the aerodynamic performance of the flapping- and revolving-wing models at their respective maximal power loadings. It is found that the revolving-wing model is more efficient than the flapping-wing model both when the body of the latter is fixed and where it can move freely. Finally, we discuss the relative agilities of the flapping- and revolving-wing models.

Vortices and turbulence (The 23rd Lanchester Memorial Lecture)

NASA Astrophysics Data System (ADS)

Lilley, G. M.

1983-12-01

A comprehensive discussion is presented concerning the phenomena characteristically treated in vortex and turbulence theory, as well as the degree of success achieved by various computation and visualization methods and theoretical models developed for vortex flow behavior prediction. Note is taken of the pioneering research conducted by F. W. Lanchester in 1893-1907, and attention is given to vortex tip and edge generation by rectangular and delta wings, the cool core effect of the Ranque-Hilsch vortex tube, the modeling of shear flows by means of vortex array methods, the classification and modelling of turbulent flows (together with a taxonomy of their calculation methods), and NASA ILLIAC IV computations of two-dimensional channel flow. Also noted are recent results concerning the boundary layer coherent structure of a flat plate at zero pressure gradient, including the regeneration structure and flow distortion and breakdown of a turbulent boundary layer.

Convair F-106B Delta Dart with Research Engines

NASA Image and Video Library

1969-08-21

A Convair F-106B Delta Dart rolls to the right to reveal the two research engines installed under its wings by the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis acquired the aircraft in October of 1966 to study inlet and nozzle designs for the supersonic transport engine program. Two General Electric J85 engines were mounted beneath the F-106B’s wings and operated from Mach 1 to 1.5. The right wing always carried reference nozzle for which the performance was known. Six supersonic nozzle variations and two inlets were tested on the left engine. The designs had already been studied in the Lewis wind tunnels, but those tests were limited by shock waves in the tunnels. Most F-106B flights were flown in a 200-mile path over the lake between Buffalo and Sandusky, known as the Lake Erie Corridor. The 1100-mile-per-hour flight took only 11 minutes at an altitude of 30,000 feet. The aircraft almost always returned with a depleted fuel supply so a Visual Flight Rules operation was required. Following the crash of another jet fighter at Lewis in July 1969, the F-106s were stationed at Selfridge Air Force Base in Michigan. NASA pilots flew transport planes each morning to the base before commencing the F-106B missions. After the supersonic transport program was cancelled, the F-106B was used as a test bed for additional engine exhaust nozzle configurations. The F-106B was also used to test inlet configurations for the noise reduction program.

Identification of aerodynamic models for maneuvering aircraft

NASA Technical Reports Server (NTRS)

Lan, C. Edward; Hu, C. C.

1992-01-01

The method based on Fourier functional analysis and indicial formulation for aerodynamic modeling as proposed by Chin and Lan is extensively examined and improved for the purpose of general applications to realistic airplane configurations. Improvement is made to automate the calculation of model coefficients, and to evaluate more accurately the indicial integral. Test data of large angle-of-attack ranges for two different models, a 70 deg. delta wing and an F-18 model, are used to further verify the applicability of Fourier functional analysis and validate the indicial formulation. The results show that the general expression for harmonic motions throughout a range of k is capable of accurately modeling the nonlinear responses with large phase lag except in the region where an inconsistent hysteresis behavior from one frequency to the other occurs. The results by the indicial formulation indicate that more accurate results can be obtained when the motion starts from a low angle of attack where hysteresis effect is not important.

Prediction of Transonic Vortex Flows Using Linear and Nonlinear Turbulent Eddy Viscosity Models

NASA Technical Reports Server (NTRS)

Bartels, Robert E.; Gatski, Thomas B.

2000-01-01

Three-dimensional transonic flow over a delta wing is investigated with a focus on the effect of transition and influence of turbulence stress anisotropies. The performance of linear eddy viscosity models and an explicit algebraic stress model is assessed at the start of vortex flow, and the results compared with experimental data. To assess the effect of transition location, computations that either fix transition or are fully turbulent are performed. To assess the effect of the turbulent stress anisotropy, comparisons are made between predictions from the algebraic stress model and the linear eddy viscosity models. Both transition location and turbulent stress anisotropy significantly affect the 3D flow field. The most significant effect is found to be the modeling of transition location. At a Mach number of 0.90, the computed solution changes character from steady to unsteady depending on transition onset. Accounting for the anisotropies in the turbulent stresses also considerably impacts the flow, most notably in the outboard region of flow separation.

Ascent heat transfer rate distribution on the North American Rockwell delta wing orbiter and the General Dynamics/Convair booster at a Mach number of 8 (mated)

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

A wind tunnel test program to determine aerodynamic interference heating on the North American Rockwell orbiter mated with the General Dynamics Convair booster is discussed. The tests were conducted at the Arnold Engineering Development Center (AEDC) in Tunnel B of the von Karman Gas Dynamics Facility (VKF). The test period was June 1971. Heat-transfer rates were determined by the phase-change paint technique on 0.013-scale Stycast models using Tempilaq as the surface temperature indicator. The nominal test conditions were: Mach 8, free-stream unit length Reynolds numbers of 1.25 x one million and 2.55 x one million angles of attack of -5, 0, +5 deg. Model details, test conditions, phase-change paint photographs and reduced heat-transfer coefficients are presented.

Static Wind-Tunnel and Radio-Controlled Flight Test Investigation of a Remotely Piloted Vehicle Having a Delta Wing Planform

NASA Technical Reports Server (NTRS)

Yip, Long P.; Fratello, David J.; Robelen, David B.; Makowiec, George M.

1990-01-01

At the request of the United States Marine Corps, an exploratory wind-tunnel and flight test investigation was conducted by the Flight Dynamics Branch at the NASA Langley Research Center to improve the stability, controllability, and general flight characteristics of the Marine Corps Exdrone RPV (Remotely Piloted Vehicle) configuration. Static wind tunnel tests were conducted in the Langley 12 foot Low Speed Wind Tunnel to identify and improve the stability and control characteristics of the vehicle. The wind tunnel test resulted in several configuration modifications which included increased elevator size, increased vertical tail size and tail moment arm, increased rudder size and aileron size, the addition of vertical wing tip fins, and the addition of leading-edge droops on the outboard wing panel to improve stall departure resistance. Flight tests of the modified configuration were conducted at the NASA Plum Tree Test Site to provide a qualitative evaluation of the flight characteristics of the modified configuration.

Vortex Flows at Supersonic Speeds

NASA Technical Reports Server (NTRS)

Wood, Richard M.; Wilcox, Floyd J., Jr.; Bauer, Steven X. S.; Allen, Jerry M.

2003-01-01

A review of research conducted at the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data are for flat plates, cavities, bodies, missiles, wings, and aircraft with Mach numbers of 1.5 to 4.6. Data are presented to show the types of vortex structures that occur at supersonic speeds and the impact of these flow structures on vehicle performance and control. The data show the presence of both small- and large-scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices. Data are shown that highlight the effect of leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber on the aerodynamics of and flow over delta wings. Finally, a discussion of a design approach for wings that use vortex flows for improved aerodynamic performance at supersonic speeds is presented.

An Overview of an Experimental Demonstration Aerotow Program

NASA Technical Reports Server (NTRS)

Murray, James E.; Bowers, Albion H.; Lokos, William A.; Peters, Todd L.; Gera, Joseph

1998-01-01

An overview of an experimental demonstration of aerotowing a delta-wing airplane with low-aspect ratio and relatively high wing loading is presented. Aerotowing of future space launch configurations is a new concept, and the objective of the work described herein is to demonstrate the aerotow operation using an airplane configuration similar to conceptual space launch vehicles. Background information on the use of aerotow for a space launch vehicle is presented, and the aerotow system used in this demonstration is described. The ground tests, analytical studies, and flight planning used to predict system behavior and to enhance flight safety are detailed. The instrumentation suite and flight test maneuvers flown are discussed, preliminary performance is assessed, and flight test results are compared with the preflight predictions.

Vortex-flow aerodynamics - An emerging design capability

NASA Technical Reports Server (NTRS)

Campbell, J. F.

1981-01-01

Promising current theoretical and simulational developments in the field of leading edge vortex-generating delta, arrow ogival wings are reported, along with the history of theory and experiment leading to them. The effects of wing slenderness, leading edge nose radius, Mach number and incidence variations, and planform on the onset of vortex generation and redistribution of aerodynamic loads are considered. The range of design possibilities in this field are consequential for the future development of strategic aircraft, supersonic transports and commercial cargo aircraft which will possess low-speed, high-lift capability by virtue of leading edge vortex generation and control without recourse to heavy and expensive leading edge high-lift devices and compound airfoils. Attention is given to interactive graphics simulation devices recently developed.

Flow field analysis of aircraft configurations using a numerical solution to the three-dimensional unified supersonic/hypersonic small disturbance equations, part 1

NASA Technical Reports Server (NTRS)

Gunness, R. C., Jr.; Knight, C. J.; Dsylva, E.

1972-01-01

The unified small disturbance equations are numerically solved using the well-known Lax-Wendroff finite difference technique. The method allows complete determination of the inviscid flow field and surface properties as long as the flow remains supersonic. Shock waves and other discontinuities are accounted for implicity in the numerical method. This technique was programed for general application to the three-dimensional case. The validity of the method is demonstrated by calculations on cones, axisymmetric bodies, lifting bodies, delta wings, and a conical wing/body combination. Part 1 contains the discussion of problem development and results of the study. Part 2 contains flow charts, subroutine descriptions, and a listing of the computer program.

Study of the vortex conditions of wings with large sweepback by extrapolation of the Jones method

NASA Technical Reports Server (NTRS)

Hirsch, P.

1980-01-01

The pockets of separation originating on the leading edges are surrounded by vortex sheets. Their configuration and intensity were determined by four conditions with the JONES approximation, which is itself corrected by a simple logic. Field pressures and stresses were computed for different cases and are compared with test results (pure deltas, swallow tails, truncations, strakes, ducks, fuselage).

Aerodynamic characteristics of the North American Rockwell space shuttle delta-wing orbiter (110C) alone and with belly-mounted external oxygen/hydrogen tanks (M equals 0.6 to 5.0)

NASA Technical Reports Server (NTRS)

Allen, E. C.

1972-01-01

Experimental aerodynamic investigations were conducted at the MSFC 14 x 14 inch trisonic wind tunnel on a .0044 scale model of the space shuttle orbiter, 110C. The test configurations included the orbiter alone and with external oxygen/hydrogen belly tanks. The six component aerodynamic force and moment data obtained cover a Mach number range from 0.6 to 4.96 at angles of attack from 10 to 60 deg at 0 deg sideslip angle and -10 to 10 deg at -6 deg sideslip angle. Reynolds number per unit length varied somewhat with Mach number but was a nominal 6.8 million per foot.

Stable isotopes document the trophic structure of a deep-sea cephalopod assemblage including giant octopod and giant squid.

PubMed

Cherel, Y; Ridoux, V; Spitz, J; Richard, P

2009-06-23

Although deep-sea cephalopods are key marine organims, their feeding ecology remains essentially unknown. Here, we report for the first time the trophic structure of an assemblage of these animals (19 species) by measuring the isotopic signature of wings of their lower beaks, which accumulated in stomachs of stranded sperm whales. Overall, the species encompassed a narrow range in delta(13)C values (1.7 per thousand), indicating that they lived in closely related and overlapping habitats. delta(13)C values can be interpreted in terms of distribution with the more (13)C-depleted species (e.g. Stigmatoteuthis arcturi, Vampyroteuthis infernalis) having a more pelagic habitat than the more (13)C-enriched, bathyal species (e.g. Todarodes sagittatus and the giant squid Architeuthis dux). The cephalopods sampled had delta(15)N values ranging 4.6 per thousand, which is consistent with the species spanning approximately 1.5 trophic levels. Neither the giant octopod (Haliphron atlanticus) nor the giant squid reached the highest trophic position. Species delta(15)N was independent of body size, with large squids having both the highest (Taningia danae) and lowest (Lepidoteuthis grimaldii) delta(15)N values. Their trophic position indicates that some species share the top of the food web, together with other megacarnivores such as the sperm whale.

The wing pattern of Moerarchis Durrant, 1914 (Lepidoptera: Tineidae) clarifies transitions between predictive models

PubMed Central

2017-01-01

The evolution of wing pattern in Lepidoptera is a popular area of inquiry but few studies have examined microlepidoptera, with fewer still focusing on intraspecific variation. The tineid genus Moerarchis Durrant, 1914 includes two species with high intraspecific variation of wing pattern. A subset of the specimens examined here provide, to my knowledge, the first examples of wing patterns that follow both the ‘alternating wing-margin’ and ‘uniform wing-margin’ models in different regions along the costa. These models can also be evaluated along the dorsum of Moerarchis, where a similar transition between the two models can be seen. Fusion of veins is shown not to effect wing pattern, in agreement with previous inferences that the plesiomorphic location of wing veins constrains the development of colour pattern. The significant correlation between wing length and number of wing pattern elements in Moerarchis australasiella shows that wing size can act as a major determinant of wing pattern complexity. Lastly, some M. australasiella specimens have wing patterns that conform entirely to the ‘uniform wing-margin’ model and contain more than six bands, providing new empirical insight into the century-old question of how wing venation constrains wing patterns with seven or more bands. PMID:28405390

The effects of wing flexibility on the flight performance and stability of flapping wing micro air vehicles

NASA Astrophysics Data System (ADS)

Bluman, James Edward

Insect wings are flexible. However, the influence of wing flexibility on the flight dynamics of insects and flapping wing micro air vehicles is unknown. Most studies in the literature consider rigid wings and conclude that the hover equilibrium is unstable. This dissertation shows that a flapping wing flyer with flexible wings exhibits stable natural modes of the open loop system in hover, never reported before. The free-flight insect flight dynamics is modeled for both flexible and rigid wings. Wing mass and inertia are included in the nonlinear equations of motion. The flapping wing aerodynamics are modeled using a quasi-steady model, a well-validated two dimensional Navier Stokes model, and a coupled, two dimensional Navier Stokes - Euler Bernoulli beam model that accurately models the fluid-structure interaction of flexible wings. Hover equilibrium is systematically and efficiently determined with a coupled quasi-steady and Navier-Stokes equation trimmer. The power and stability are reported at hover while parametrically varying the pitch axis location for rigid wings and the structural stiffness for flexible wings. The results indicate that the rigid wings possess an unstable oscillatory mode mainly due to their pitch sensitivity to horizontal velocity perturbations. The flexible wings stabilize this mode primarily by adjusting their wing shape in the presence of perturbations. The wing's response to perturbations generates significantly more horizontal velocity damping and pitch rate damping than in rigid wings. Furthermore, the flexible wings experience substantially less wing wake interaction, which, for rigid wings, is destabilizing. The power required to hover a fruit fly with actively rotating rigid wings varies between 16.9 and 34.2 W/kg. The optimal power occurs when the pitch axis is located at 30% chord, similar to some biological observations. Flexible wings require 23.1 to 38.5 W/kg. However, flexible wings exhibit more stable system dynamics and allow for simpler and lighter designs since they do not require pitch actuation mechanisms. This study is the first to evaluate the impact of wing flexibility on the hovering stability of flapping flyers, which can explain the ranges of flexibility seen in insects and can inform designs of synthetic flapping wing robots.

Plasma Control of Separated Flows on Delta Wings at High Angles of Attack

DTIC Science & Technology

2009-03-18

of Attack 5a. CONTRACT NUMBER ISTC Registration No: 3646 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) Dr. Anatoly Alexandrovich...NUMBER(S) ISTC 06-7002 12. DISTRIBUTION/AVAILABILITY STATEMENT Approved for public release; distribution is unlimited. 13. SUPPLEMENTARY...This work is supported financially by EOARD and performed under the agreement with the International Science and Technology Center ( ISTC ), Moscow

Aircraft integrated design and analysis: A classroom experience

NASA Technical Reports Server (NTRS)

Weisshaar, Terrence A.

1989-01-01

AAE 451 is the capstone course required of all senior undergraduates in the School of Aeronautics and Astronautics at Purdue University. During the past year the first steps of a long evolutionary process were taken to change the content and expectations of this course. These changes are the result of the availability of advanced computational capabilities and sophisticated electronic media availability at Purdue. This presentation will describe both the long range objectives and this year's experience using the High Speed Commercial Transport design, the AIAA Long Duration Aircraft design and RPV design proposal as project objectives. The central goal of these efforts is to provide a user-friendly, computer-software-based environment to supplement traditional design course methodology. The Purdue University Computer Center (PUCC), the Engineering Computer Network (ECN) and stand-alone PC's are being used for this development. This year's accomplishments center primarily on aerodynamics software obtained from NASA/Langley and its integration into the classroom. Word processor capability for oral and written work and computer graphics were also blended into the course. A total of ten HSCT designs were generated, ranging from twin-fuselage aircraft, forward swept wing aircraft to the more traditional delta and double-delta wing aircraft. Four Long Duration Aircraft designs were submitted, together with one RPV design tailored for photographic surveillance.

Independent Orbiter Assessment (IOA): Analysis of the elevon subsystem

NASA Technical Reports Server (NTRS)

Wilson, R. E.; Riccio, J. R.

1986-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results for the Orbiter Elevon system hardware. The elevon actuators are located at the trailing edge of the wing surface. The proper function of the elevons is essential during the dynamic flight phases of ascent and entry. In the ascent phase of flight, the elevons are used for relieving high wing loads. For entry, the elevons are used to pitch and roll the vehicle. Specifically, the elevon system hardware comprises the following components: flow cutoff valve; switching valve; electro-hydraulic (EH) servoactuator; secondary delta pressure transducer; bypass valve; power valve; power valve check valve; primary actuator; primary delta pressure transducer; and primary actuator position transducer. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Of the 25 failure modes analyzed, 18 were determined to be PCIs.

Subsonic balance and pressure investigation of a 60 deg delta wing with leading edge devices

NASA Technical Reports Server (NTRS)

Tingas, S. A.; Rao, D. M.

1982-01-01

Low supersonic wave drag makes the thin highly swept delta wing the logical choice for use on aircraft designed for supersonic cruise. However, the high-lift maneuver capability of the aircraft is limited by severe induced-drag penalties attributed to loss of potential flow leading-edge suction. This drag increase may be alleviated through leading-edge flow control to recover lost aerodynamic thrust through either retention of attached leading-edge flow to higher angles of attack or exploitation of the increased suction potential of separation-induced vortex flow. A low-speed wind-tunnel investigation was undertaken to examine the high-lift devices such as fences, chordwise slots, pylon vortex generators, leading-edge vortex flaps, and sharp leading-edge extensions. The devices were tested individually and in combinations in an attempt to improve high-alpha drag performance with a minimum of low-alpha drag penalty. This report presents an analysis of the force, moment, and static pressure data obtained in angles of attack up to 23 deg, at Mach and Reynolds numbers of 0.16 and 3.85 x 10 to the 6th power per meter, respectively. The results indicate that all the devices produced drag and longitudinal/lateral stability improvements at high lift with, in most cases, minor drag penalties at low angles of attack.

Laboratory modeling and analysis of aircraft-lightning interactions

NASA Technical Reports Server (NTRS)

Turner, C. D.; Trost, T. F.

1982-01-01

Modeling studies of the interaction of a delta wing aircraft with direct lightning strikes were carried out using an approximate scale model of an F-106B. The model, which is three feet in length, is subjected to direct injection of fast current pulses supplied by wires, which simulate the lightning channel and are attached at various locations on the model. Measurements are made of the resulting transient electromagnetic fields using time derivative sensors. The sensor outputs are sampled and digitized by computer. The noise level is reduced by averaging the sensor output from ten input pulses at each sample time. Computer analysis of the measured fields includes Fourier transformation and the computation of transfer functions for the model. Prony analysis is also used to determine the natural frequencies of the model. Comparisons of model natural frequencies extracted by Prony analysis with those for in flight direct strike data usually show lower damping in the in flight case. This is indicative of either a lightning channel with a higher impedance than the wires on the model, only one attachment point, or short streamers instead of a long channel.

Wind tunnel tests of a free-wing/free-trimmer model

NASA Technical Reports Server (NTRS)

Sandlin, D. R.

1982-01-01

The riding qualities of an aircraft with low wing loading can be improved by freeing the wing to rotate about its spanwise axis. A trimming surface also free to rotate about its spanwise axis can be added at the wing tips to permit the use of high lift devices. Wind tunnel tests of the free wing/free trimmer model with the trimmer attached to the wing tips aft of the wing chord were conducted to validate a mathematical model developed to predict the dynamic characteristics of a free wing/free trimmer aircraft. A model consisting of a semispan wing with the trimmer mounted on with the wing on an air bearing and the trimmer on a ball bearing was displaced to various angles of attack and released. The damped oscillations of the wing and trimmer were recorded. Real and imaginary parts of the characteristic equations of motion were determined and compared to values predicted using the mathematical model.

Heat transfer rate distribution on North American Rockwell delta wing orbiter determined by phase change paint technique at a Mach number of 8, volume 1

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

The results of a wind tunnel test program to determine aerodynamic heat transfer distributions on an orbiter configuration are presented. Heat-transfer rates were determined by the phase change paint technique on 0.013-scale Stycast models using Tempilaq as the surface temperature indicator. The nominal test conditions were; Mach 8, length Reynolds numbers of 6.0 x 1 million and 8.9 x 1 million, and angles of attack from 10 to 50 deg in 10-deg increments. At the higher Reynolds number, data were obtained with and without boundary layer trips. Model details, test conditions, and reduced heat-transfer data are presented. Data reduction of the phase-change paint photographs was performed by utilizing a new technique which is described in the data presentation section.

The Effects of a Highly Cambered Low-Drag Wing and of Auxiliary Flaps on the High-Speed Aerodynamic Characteristics of a Twin-Engine Pursuit Airplane Model

NASA Technical Reports Server (NTRS)

Ganzer, Victor M

1944-01-01

Results are presented for tests of two wings, an NACA 230-series wing and a highly-cambered NACA 66-series wing on a twin-engine pursuit airplane. Auxiliary control flaps were tested in combinations with each wing. Data showing comparison of high-speed aerodynamic characteristics of the model when equipped with each wing, the effect of the auxiliary control flaps on aerodynamic characteristics, and elevator effectiveness for the model with the 66-series wing are presented. High-speed aerodynamic characteristics of the model were improved with the 66-series wing.

KSC-00padig074

NASA Image and Video Library

2000-11-02

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, participants cut the lines to helium-filled balloons. From left, they are Center Director Roy Bridges; Michael Butchko, president, SGS; Pierre Dufour, president and CEO, Air Liquide America Corporation; David Herst, director, Delta IV Launch Sites; Pamela Gillespie, executive administrator, office of Congressman Dave Weldon; and Col. Samuel Dick, representative of the 45th Space Wing. The nine-mile-long buried pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. It will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS), and Ramon Lugo, acting executive director, JPMO

KSC-00padig075

NASA Image and Video Library

2000-11-02

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, participants watch as helium-filled balloons take to the sky after their lines were cut. From left, they are Center Director Roy Bridges; Michael Butchko, president, SGS; Pierre Dufour, president and CEO, Air Liquide America Corporation; David Herst, director, Delta IV Launch Sites; Pamela Gillespie, executive administrator, office of Congressman Dave Weldon; and Col. Samuel Dick, representative of the 45th Space Wing. The nine-mile-long buried pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. It will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS), and Ramon Lugo, acting executive director, JPMO

KSC00padig075

NASA Image and Video Library

2000-11-02

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, participants watch as helium-filled balloons take to the sky after their lines were cut. From left, they are Center Director Roy Bridges; Michael Butchko, president, SGS; Pierre Dufour, president and CEO, Air Liquide America Corporation; David Herst, director, Delta IV Launch Sites; Pamela Gillespie, executive administrator, office of Congressman Dave Weldon; and Col. Samuel Dick, representative of the 45th Space Wing. The nine-mile-long buried pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. It will also serve as a backup helium resource for Shuttle launches. Nearly one launch’s worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS), and Ramon Lugo, acting executive director, JPMO

Commissioning of a new helium pipeline

NASA Technical Reports Server (NTRS)

2000-01-01

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, participants cut the lines to helium-filled balloons. From left, they are Center Director Roy Bridges; Michael Butchko, president, SGS; Pierre Dufour, president and CEO, Air Liquide America Corporation; David Herst, director, Delta IV Launch Sites; Pamela Gillespie, executive administrator, office of Congressman Dave Weldon; and Col. Samuel Dick, representative of the 45th Space Wing. The nine-mile-long buried pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. It will also serve as a backup helium resource for Shuttle launches. Nearly one launch's worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS), and Ramon Lugo, acting executive director, JPMO.

Commissioning of a new helium pipeline

NASA Technical Reports Server (NTRS)

2000-01-01

At the commissioning of a new high-pressure helium pipeline at Kennedy Space Center, participants watch as helium-filled balloons take to the sky after their lines were cut. From left, they are Center Director Roy Bridges; Michael Butchko, president, SGS; Pierre Dufour, president and CEO, Air Liquide America Corporation; David Herst, director, Delta IV Launch Sites; Pamela Gillespie, executive administrator, office of Congressman Dave Weldon; and Col. Samuel Dick, representative of the 45th Space Wing. The nine-mile-long buried pipeline will service launch needs at the new Delta IV Complex 37 at Cape Canaveral Air Force Station. It will also serve as a backup helium resource for Shuttle launches. Nearly one launch's worth of helium will be available in the pipeline to support a Shuttle pad in an emergency. The line originates at the Helium Facility on KSC and terminates in a meter station at the perimeter of the Delta IV launch pad. Others at the ceremony were Jerry Jorgensen, pipeline project manager, Space Gateway Support (SGS), and Ramon Lugo, acting executive director, JPMO.

Dual methods and approximation concepts in structural synthesis

NASA Technical Reports Server (NTRS)

Fleury, C.; Schmit, L. A., Jr.

1980-01-01

Approximation concepts and dual method algorithms are combined to create a method for minimum weight design of structural systems. Approximation concepts convert the basic mathematical programming statement of the structural synthesis problem into a sequence of explicit primal problems of separable form. These problems are solved by constructing explicit dual functions, which are maximized subject to nonnegativity constraints on the dual variables. It is shown that the joining together of approximation concepts and dual methods can be viewed as a generalized optimality criteria approach. The dual method is successfully extended to deal with pure discrete and mixed continuous-discrete design variable problems. The power of the method presented is illustrated with numerical results for example problems, including a metallic swept wing and a thin delta wing with fiber composite skins.

Investigation of safe-life fail-safe criteria for the space shuttle

NASA Technical Reports Server (NTRS)

1972-01-01

An investigation was made to determine the effects of a safe-life design approach and a fail-safe design approach on the space shuttle booster vehicle structure, and to recommend any changes to the structural design criteria. Two configurations of the booster vehicle were considered, one incorporating a delta wing (B-9U configuration) and the other a swept wing (B-16B configuration). Several major structural components of the booster were studied to determine the fatigue life, safe-life, and fail-safe capabilities of the baseline design. Each component was investigated to determine the practicability of applying a safe-life or fail-safe design philosophy, the changes such design approaches might require, and the impact of these changes on weight, cost, development plans, and performance.

Study of aerodynamic technology for VSTOL fighter/attack aircraft: Vertical attitude concept

NASA Technical Reports Server (NTRS)

Gerhardt, H. A.; Chen, W. S.

1978-01-01

The aerodynamic technology for a vertical attitude VSTOL (VATOL) supersonic fighter/attack aircraft was studied. The selected configuration features a tailless clipped delta wing with leading-edge extension (LEX), maneuvering flaps, top-side inlet, twin dry engines and vectoring nozzles. A relaxed static stability is employed in conjunction with the maneuvering flaps to optimize transonic performance and minimize supersonic trim drag. Control for subaerodynamic flight is obtained by gimballing the nozzles in combination with wing tip jets. Emphasis is placed on the development of aerodynamic characteristics and the identification of aerodynamic uncertainties. A wind tunnel test program is proposed to resolve these uncertainties and ascertain the feasibility of the conceptual design. Ship interface, flight control integration, crew station concepts, advanced weapons, avionics, and materials are discussed.

Active vibration suppression of self-excited structures using an adaptive LMS algorithm

NASA Astrophysics Data System (ADS)

Danda Roy, Indranil

The purpose of this investigation is to study the feasibility of an adaptive feedforward controller for active flutter suppression in representative linear wing models. The ability of the controller to suppress limit-cycle oscillations in wing models having root springs with freeplay nonlinearities has also been studied. For the purposes of numerical simulation, mathematical models of a rigid and a flexible wing structure have been developed. The rigid wing model is represented by a simple three-degree-of-freedom airfoil while the flexible wing is modelled by a multi-degree-of-freedom finite element representation with beam elements for bending and rod elements for torsion. Control action is provided by one or more flaps attached to the trailing edge and extending along the entire wing span for the rigid model and a fraction of the wing span for the flexible model. Both two-dimensional quasi-steady aerodynamics and time-domain unsteady aerodynamics have been used to generate the airforces in the wing models. An adaptive feedforward controller has been designed based on the filtered-X Least Mean Squares (LMS) algorithm. The control configuration for the rigid wing model is single-input single-output (SISO) while both SISO and multi-input multi-output (MIMO) configurations have been applied on the flexible wing model. The controller includes an on-line adaptive system identification scheme which provides the LMS controller with a reasonably accurate model of the plant. This enables the adaptive controller to track time-varying parameters in the plant and provide effective control. The wing models in closed-loop exhibit highly damped responses at airspeeds where the open-loop responses are destructive. Simulations with the rigid and the flexible wing models in a time-varying airstream show a 63% and 53% increase, respectively, over their corresponding open-loop flutter airspeeds. The ability of the LMS controller to suppress wing store flutter in the two models has also been investigated. With 10% measurement noise introduced in the flexible wing model, the controller demonstrated good robustness to the extraneous disturbances. In the examples studied it is found that adaptation is rapid enough to successfully control flutter at accelerations in the airstream of up to 15 ft/sec2 for the rigid wing model and 9 ft/sec2 for the flexible wing model.

Flexible Wing Model for Structural Sizing and Multidisciplinary Design Optimization of a Strut-Braced Wing

NASA Technical Reports Server (NTRS)

Gern, Frank H.; Naghshineh, Amir H.; Sulaeman, Erwin; Kapania, Rakesh K.; Haftka, Raphael T.

2000-01-01

This paper describes a structural and aeroelastic model for wing sizing and weight calculation of a strut-braced wing. The wing weight is calculated using a newly developed structural weight analysis module considering the special nature of strut-braced wings. A specially developed aeroelastic model enables one to consider wing flexibility and spanload redistribution during in-flight maneuvers. The structural model uses a hexagonal wing-box featuring skin panels, stringers, and spar caps, whereas the aerodynamics part employs a linearized transonic vortex lattice method. Thus, the wing weight may be calculated from the rigid or flexible wing spanload. The calculations reveal the significant influence of the strut on the bending material weight of the wing. The use of a strut enables one to design a wing with thin airfoils without weight penalty. The strut also influences wing spanload and deformations. Weight savings are not only possible by calculation and iterative resizing of the wing structure according to the actual design loads. Moreover, as an advantage over the cantilever wing, employment of the strut twist moment for further load alleviation leads to increased savings in structural weight.

General Electric 32-Spoke Nozzle on the Convair F-106B Delta Dart

NASA Image and Video Library

1971-03-21

National Aeronautics and Space Administration (NASA) Convair F-106B Delta Dart with a 32-spoke nozzle installed on its General Electric J85 test engine. Lewis acquired a Delta Dart fighter in 1966 to study the components for propulsion systems that could be applied to supersonic transport aircraft at transonic speeds. The F-106B was modified with two General Electric J85-13 engines under its wings to study these components. The original test plan was expanded to include the study of boattail drag, noise reduction, and inlets. From February to July 1971 the modified F-106B was used to study different ejector nozzles. Researchers conducted both acoustic and aerodynamic tests on the ground and in flight. Several models were created to test different suppression methods. NASA Lewis’ conical nozzle was used as the baseline configuration. Flightline and sideline microphones were set up on the ground. The F-106B would idle its own engine and buzz the recording station from an altitude of 300 feet at Mach 0.4 with the test engines firing. Researchers found that the suppression of the perceived noise level was usually lower during flight than the researchers had statistically predicted. The 64 and 32-spoke nozzles performed well in actual flight, but the others nozzles tended to negatively affect the engine’s performance. Different speeds or angles- -of-attack sometimes changed the noise levels. In the end, no general conclusions could be applied to all the nozzles.

Influence of numerical dissipation in computing supersonic vortex-dominated flows

NASA Technical Reports Server (NTRS)

Kandil, O. A.; Chuang, A.

1986-01-01

Steady supersonic vortex-dominated flows are solved using the unsteady Euler equations for conical and three-dimensional flows around sharp- and round-edged delta wings. The computational method is a finite-volume scheme which uses a four-stage Runge-Kutta time stepping with explicit second- and fourth-order dissipation terms. The grid is generated by a modified Joukowski transformation. The steady flow solution is obtained through time-stepping with initial conditions corresponding to the freestream conditions, and the bow shock is captured as a part of the solution. The scheme is applied to flat-plate and elliptic-section wings with a leading edge sweep of 70 deg at an angle of attack of 10 deg and a freestream Mach number of 2.0. Three grid sizes of 29 x 39, 65 x 65 and 100 x 100 have been used. The results for sharp-edged wings show that they are consistent with all grid sizes and variation of the artificial viscosity coefficients. The results for round-edged wings show that separated and attached flow solutions can be obtained by varying the artificial viscosity coefficients. They also show that the solutions are independent of the way time stepping is done. Local time-stepping and global minimum time-steeping produce same solutions.

SMA actuators for morphing wings

NASA Astrophysics Data System (ADS)

Brailovski, V.; Terriault, P.; Georges, T.; Coutu, D.

An experimental morphing laminar wing was developed to prove the feasibility of aircraft fuel consumption reduction through enhancement of the laminar flow regime over the wing extrados. The morphing wing prototype designed for subsonic cruise flight conditions (Mach 0.2 … 0.3; angle of attack - 1 … +2∘), combines three principal subsystems: (1) flexible extrados, (2) rigid intrados and (3) an actuator group located inside the wing box. The morphing capability of the wing relies on controlled deformation of the wing extrados under the action of shape memory alloys (SMA) actuators. A coupled fluid-structure model of the morphing wing was used to evaluate its mechanical and aerodynamic performances in different flight conditions. A 0.5 m chord and 1 m span prototype of the morphing wing was tested in a subsonic wind tunnel. In this work, SMA actuators for morphing wings were modeled using a coupled thermo-mechanical finite element model and they were windtunnel validated. If the thermo-mechanical model of SMA actuators presented in this work is coupled with the previously developed structureaerodynamic model of the morphing wing, it could serve for the optimization of the entire morphing wing system.

An experimental analysis of critical factors involved in the breakdown process of leading edge vortex flows. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Visser, Kenneth D.

1991-01-01

Experimental crosswire measurements of the flowfield above a 70 and 75 degree flat plate delta wing were performed at a Reynolds number of 250,000. Survey grids were taken normal to the platform at a series of chordwise locations for angles of attack of 20 and 30 degrees. Axial and azimuthal vorticity distributions were derived from the velocity fields. The dependence of circulation on distance from the vortex core as well as on chordwise location was examined. The effects of nondimensionalization in comparison with other experimental data was made. The circulation distribution scales with the local semispan and grows approximately linearly in the chordwise direction. For regions of the flow outside of the vortex subcore, the circulation at any chordwise station was observed to vary logarithmically with distance from the vortex axis. The circulation was also found to increase linearly with angle of incidence at a given chordwise station. A reduction in the local circulation about the vortex axis occurred at breakdown. The spanwise distribution of axial vorticity was severely altered through the breakdown region and the spanwise distribution of axial vorticity present appeared to reach a maximum immediately preceding breakdown. The local concentration of axial vorticity about the vortex axis was reduced while the magnitude of the azimuthal vorticity decreased throughout the breakdown zone. The axial vorticity components with a negative sense, found in the secondary vortex, remained unaffected by changes in wing sweep or angle of attack, in direct contrast to the positive components. The inclusion of the local wing geometry into a previously derived correlation parameter indicated that the circulation of growing leading edge vortex flows were similar at corresponding radii from the vortex axis. It was concluded that the flow over a delta wing, upstream of the breakdown regions and away from the apex and trailing edge regions, is conical. In addition, the dominating factors leading to the onset of breakdown are felt to be the local circulation of the vortex and the accompanying pressure field.

F-16XL Ship #2 wing glove close-up, laser cut holes, with dime for scale

NASA Technical Reports Server (NTRS)

1995-01-01

This June 1995 photograph of a test panel similiar to the one attached to the surface of an F-16XL research aircraft's left wing at NASA's Dryden Flight Research Center, Edwards, California, shows the size of the more than 10 million laser-cut holes in the panel, called a glove, as compared with a dime. Below the titanium panel into which holes are cut is a suction system linked to a compressor. During research flights with the modified, delta-winged F-16XL, the suction system pulled a small part of the boundary layer of air through the glove's porous surface to expand the extent of smooth (laminar) flow. Researchers believe that laminar flow conditions can reduce aerodynamic drag (friction) and contribute to reduced operating costs by improving fuel consumption and lowering aircraft weight. This Supersonic Laminar Flow Control (SLFC) experiment represents a collaborative effort between NASA and aerospace industry (specifically Boeing, Rockwell, and McDonnell Douglas), with Boeing assembling the panel and McDonnell Douglas designing the suction system.

Pegasus(Registered trademark) Wing-Glove Experiment to Document Hypersonic Crossflow Transition: Measurement System and Selected Flight Results

NASA Technical Reports Server (NTRS)

Bertelrud, Arild; delaTova, Geva; Hamory, Philip J.; Young, Ronald; Noffz, Gregory K.; Dodson, Michael; Graves, Sharon S.; Diamond, John K.; Bartlett, James E.; Noack, Robert;

Parachuting with bristled wings

NASA Astrophysics Data System (ADS)

Kasoju, Vishwa; Santhanakrishnan, Arvind; Senter, Michael; Armel, Kristen; Miller, Laura

2017-11-01

Free takeoff flight recordings of thrips (body length <1 mm) show that they can intermittently cease flapping and instead float passively downwards by spreading their bristled wings. Such drag-based parachuting can lower the speed of falling and aid in long distance dispersal by minimizing energetic demands needed for active flapping flight. However, the role of bristled wings in parachuting remains unclear. In this study, we examine if using bristled wings lowers drag forces in parachuting as compared to solid (non-bristled) wings. Wing angles and settling velocities were obtained from free takeoff flight videos. A solid wing model and bristled wing model with bristle spacing to diameter ratio of 5 performing translational motion were comparatively examined using a dynamically scaled robotic model. We measured force generated under varying wing angle from 45-75 degrees across a Reynolds number (Re) range of 1 to 15. Drag experienced by the wings decreased in both wing models when varying Re from 1 to 15. Leakiness of flow through bristles, visualized using spanwise PIV, and implications for force generation will be presented. Numerical simulations will be used to investigate the stability of free fall using bristled wings.

A qualitative study of vortex trapping capability for lift enhancement on unconventional wing

NASA Astrophysics Data System (ADS)

Salleh, M. B.; Kamaruddin, N. M.; Mohamed-Kassim, Z.

2018-05-01

Lift enhancement by using passive vortex trapping technique offers great advantage in small aircraft design as it can improve aerodynamics performance and reduce weight of the wing. To achieve this aim, a qualitative study on the flow structures across wing models with cavities has been performed using smoke wire visualisation technique. An experiment has been conducted at low Reynolds number of 26,000 with angle of attack (α) = 0°, 5°, 10° and 15° to investigate the vortex trapping capability of semi-circular leading edge (SCLE) flat-plate wing model and elliptical leading edge (ELE) flat-plate wing model with cavities, respectively. Results from the qualitative study indicated unique characteristics in the flow structures between the tested wing models. The SCLE wing models were able to trap stable rotating vortices for α ≤ 10° whereas the ability of ELE wing models to suppress flow separation allowed stable clockwise vortices to be trapped inside the cavities even at α > 10°. The trapped vortices found to have the potential to increase lift on the unconventional wing models.

An integrated CFD/experimental analysis of aerodynamic forces and moments

NASA Technical Reports Server (NTRS)

Melton, John E.; Robertson, David D.; Moyer, Seth A.

1989-01-01

Aerodynamic analysis using computational fluid dynamics (CFD) is most fruitful when it is combined with a thorough program of wind tunnel testing. The understanding of aerodynamic phenomena is enhanced by the synergistic use of both analysis methods. A technique is described for an integrated approach to determining the forces and moments acting on a wind tunnel model by using a combination of experimentally measured pressures and CFD predictions. The CFD code used was FLO57 (an Euler solver) and the wind tunnel model was a heavily instrumented delta wing with 62.5 deg of leading-edge sweep. A thorough comparison of the CFD results and the experimental data is presented for surface pressure distributions and longitudinal forces and moments. The experimental pressures were also integrated over the surface of the model and the resulting forces and moments are compared to the CFD and wind tunnel results. The accurate determination of various drag increments via the combined use of the CFD and experimental pressures is presented in detail.

KSC-2014-1989

NASA Image and Video Library

2014-03-20

VANDENBERG AIR FORCE BASE, Calif. – The Delta first-stage booster for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, makes a steep turn toward the Horizontal Processing Facility at Space Launch Complex 2 on Vandenberg Air Force Base in California. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket on July 1, 2014. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/D. Liberotti, 30th Space Wing, VAFB

KSC-2014-1992

NASA Image and Video Library

2014-03-20

VANDENBERG AIR FORCE BASE, Calif. – The Delta first-stage booster for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, rests on its hardware transportation cradle in the Horizontal Processing Facility at Space Launch Complex 2 on Vandenberg Air Force Base in California. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket on July 1, 2014. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/D. Liberotti, 30th Space Wing, VAFB

How wing compliance drives the efficiency of self-propelled flapping flyers.

PubMed

Thiria, Benjamin; Godoy-Diana, Ramiro

2010-07-01

Wing flexibility governs the flying performance of flapping-wing flyers. Here, we use a self-propelled flapping-wing model mounted on a "merry go round" to investigate the effect of wing compliance on the propulsive efficiency of the system. Our measurements show that the elastic nature of the wings can lead not only to a substantial reduction in the consumed power, but also to an increment of the propulsive force. A scaling analysis using a flexible plate model for the wings points out that, for flapping flyers in air, the time-dependent shape of the elastic bending wing is governed by the wing inertia. Based on this prediction, we define the ratio of the inertial forces deforming the wing to the elastic restoring force that limits the deformation as the elastoinertial number N(ei). Our measurements with the self-propelled model confirm that it is the appropriate structural parameter to describe flapping flyers with flexible wings.

Effect of wing mass in free flight by a butterfly-like 3D flapping wing-body model

NASA Astrophysics Data System (ADS)

Suzuki, Kosuke; Okada, Iori; Yoshino, Masato

2016-11-01

The effect of wing mass in free flight of a flapping wing is investigated by numerical simulations based on an immersed boundary-lattice Boltzmann method. We consider a butterfly-like 3D flapping wing-model consisting of two square wings with uniform mass density connected by a rod-shaped body. We simulate free flights of the wing-body model with various mass ratios of the wing to the whole of the model. As a result, it is found that the lift and thrust forces decrease as the mass ratio increases, since the body with a large mass ratio experiences large vertical and horizontal oscillations in one period and consequently the wing tip speed relatively decreases. In addition, we find the critical mass ratio between upward flight and downward flight for various Reynolds numbers. This work was supported by JSPS KAKENHI Grant Number JP16K18012.

The timing of wing molt in tundra swans: energetic and non-energetic constraints

USGS Publications Warehouse

Earnst, S.L.

1992-01-01

Date of wing molt initiation, based on the regression of tenth primary length on capture date, was calculated for breeding and nonbreeding Tundra Swans (Cygnus columbianus columbianus) on the Colville River Delta, Alaska. Breeding females initiated wing molt significantly later than breeding males and nonbreeding males and females; the molt of breeding females was correlated with the date on which their eggs hatched. Breeding males did not differ significantly from nonbreeding males and females in the date of molt initiation. Timing of molt in breeding males and females was consistent with the views that females delay molt while replenishing energy spent on reproduction, but was also consistent with the breeding pair's need for primaries to defend territories and to defend and brood young. Other results, including an increase in an index of female body condition throughout most of the molt period, and a positive correlation between clutch size and female hatch-to-molt interval, were not predicted by the hypothesis that past energy expenditures constrain the timing of molt. Patterns of wing molt within and among other Northern Hemisphere geese and swans are also difficult to explain on the basis of energetics alone. For example, breeding females initiate molt before breeding males in many species. Also, there is extreme asynchrony between mates in two swan species; one of those species also exhibits variation in which sex initiates wing molt first. Both patterns suggest that asynchrony, per se, is important, probably to facilitate brood protection or territory defense. In Tundra Swans and other northern breeding geese and swans, the non-energetic demands of territory defense, brood defense, and brooding are probably important constraints on the timing of wing molt.

MULTISHOCKED,THREE-DIMENSIONAL SUPERSONIC FLOWFIELDS WITH REAL GAS EFFECTS

NASA Technical Reports Server (NTRS)

Kutler, P.

1994-01-01

This program determines the supersonic flowfield surrounding three-dimensional wing-body configurations of a delta wing. It was designed to provide the numerical computation of three dimensional inviscid, flowfields of either perfect or real gases about supersonic or hypersonic airplanes. The governing equations in conservation law form are solved by a finite difference method using a second order noncentered algorithm between the body and the outermost shock wave, which is treated as a sharp discontinuity. Secondary shocks which form between these boundaries are captured automatically. The flowfield between the body and outermost shock is treated in a shock capturing fashion and therefore allows for the correct formation of secondary internal shocks . The program operates in batch mode, is in CDC update format, has been implemented on the CDC 7600, and requires more than 140K (octal) word locations.

Longitudinal Stability and Control Characteristics from a Flight Investigation of a Cruciform Canard Missile Configuration Having an Exposed Wing-canard Area Ratio of 16:1

NASA Technical Reports Server (NTRS)

Moul, Martin T; Wineman, Andrew R

1952-01-01

A flight investigation has been made to determine the longitudinal stability and control characteristics of a 60 0 delta-wing-canard missile configuration with an exposed wing-canard area ratio of 16:1. The results presented include the longitudinal stability derivatives, control effectiveness, and drag characteristics for a Mach number range of 0.75 to 1.80 and are compared with the results of a similar configuration having larger 6ontrols. Stability characteristics are also presented from the flights of an interdigitated canard configuration at a Mach number of 2.08 and a wing-body configuration at Mach numbers of 1.25 to 1.45. The stability derivatives varied gradually with Mach number with the exception of the damping-in-pitch derivative. Aerodynamic damping in pitch decreased to a minimum at a Mach number of 1.0 3, then increased to a peak value at a Mach number of 1.26 followed by a gradual decrease at higher Mach numbers. The aerodynamic-center location of the in-line canard configuration shifted rearward 13 percent of the mean aerodynamic chord at transonic speeds. The pitching-moment curve slope was 25 percent greater for the model having no canards than for the in-line configuration. No large effects of interdigitation were noted in the stability derivatives. Pitching effectiveness of the in-line configuration was maintained throughout the Mach number range. A comparison of the stability and control characteristics of two canard configurations having different area controls showed that decreasing the control area 44 percent decreased the pitching effectiveness proportionally, shifted the aerodynamic-center location rearward 9 to 14 percent of the mean aerodynamic chord, and reduced the total hinge moments required for 10 trimmed flight about 50 percent at transonic speeds.

Uncertainty Quantification and Certification Prediction of Low-Boom Supersonic Aircraft Configurations

NASA Technical Reports Server (NTRS)

West, Thomas K., IV; Reuter, Bryan W.; Walker, Eric L.; Kleb, Bil; Park, Michael A.

2014-01-01

The primary objective of this work was to develop and demonstrate a process for accurate and efficient uncertainty quantification and certification prediction of low-boom, supersonic, transport aircraft. High-fidelity computational fluid dynamics models of multiple low-boom configurations were investigated including the Lockheed Martin SEEB-ALR body of revolution, the NASA 69 Delta Wing, and the Lockheed Martin 1021-01 configuration. A nonintrusive polynomial chaos surrogate modeling approach was used for reduced computational cost of propagating mixed, inherent (aleatory) and model-form (epistemic) uncertainty from both the computation fluid dynamics model and the near-field to ground level propagation model. A methodology has also been introduced to quantify the plausibility of a design to pass a certification under uncertainty. Results of this study include the analysis of each of the three configurations of interest under inviscid and fully turbulent flow assumptions. A comparison of the uncertainty outputs and sensitivity analyses between the configurations is also given. The results of this study illustrate the flexibility and robustness of the developed framework as a tool for uncertainty quantification and certification prediction of low-boom, supersonic aircraft.

An Evaluation of the Argentinean Basic Trainer Aircraft Domestic Development Project

DTIC Science & Technology

2012-03-01

Prototype, 1st jet built in Latin America 1947 IAe 31 Colibrí Two- seat Trainer aircraft 3 National design 1948 IAe 30 Ñancú Fighter/Attack prototype...37 Supersonic delta-wing interceptor (Glider, unpowered prototype only) 1 Designed by Reimar Horten. 1957 IAe 46 Ranquel 2- seat utility...return all surfaces to neutral. It must be operable from both positions, with priority on the rear command seat . • Ergonomic Throttle controls on the

Surface pressure and inviscid flow field properties of the McDonnell-Douglas delta-wing orbiter for nominal Mach number of 8, Volume 1

NASA Technical Reports Server (NTRS)

Warmbrod, J. D.; Martindale, M. R.; Matthews, R. K.

1972-01-01

The results of a wind tunnel test program to determine the surface pressures and flow distribution on the McDonnell Douglas Orbiter configuration are presented. Tests were conducted in hypersonic wind tunnel at Mach 8. The freestream unit Reynolds number was 3.7 time one million per foot. Angle of attack was varied from 10 degrees to 60 degrees in 10 degree increments.

The design of two-stage-to-orbit vehicles

NASA Technical Reports Server (NTRS)

1991-01-01

Two separate student design groups developed conceptual designs for a two-stage-to-orbit vehicle, with each design group consisting of a carrier team and an orbiter team. A two-stage-to-orbit system is considered in the event that single-stage-to-orbit is deemed not feasible in the foreseeable future; the two-stage system would also be used as a complement to an already existing heavy lift vehicle. The design specifications given are to lift a 10,000-lb payload 27 ft long by 10 ft diameter, to low Earth orbit (300 n.m.) using an air breathing carrier configuration that will take off horizontally within 15,000 ft. The staging Mach number and altitude were to be determined by the design groups. One group designed a delta wing/body carrier with the orbiter nested within the fuselage of the carrier, and the other group produced a blended cranked-delta wing/body carrier with the orbiter in the more conventional piggyback configuration. Each carrier used liquid hydrogen-fueled turbofanramjet engines, with data provided by General Electric Aircraft Engine Group. While one orbiter used a full-scale Space Shuttle Main Engine (SSME), the other orbiter employed a half-scale SSME coupled with scramjet engines, with data again provided by General Electric. The two groups conceptual designs, along with the technical trade-offs, difficulties, and details that surfaced during the design process are presented.

Static Aeroelastic Scaling and Analysis of a Sub-Scale Flexible Wing Wind Tunnel Model

NASA Technical Reports Server (NTRS)

Ting, Eric; Lebofsky, Sonia; Nguyen, Nhan; Trinh, Khanh

2014-01-01

This paper presents an approach to the development of a scaled wind tunnel model for static aeroelastic similarity with a full-scale wing model. The full-scale aircraft model is based on the NASA Generic Transport Model (GTM) with flexible wing structures referred to as the Elastically Shaped Aircraft Concept (ESAC). The baseline stiffness of the ESAC wing represents a conventionally stiff wing model. Static aeroelastic scaling is conducted on the stiff wing configuration to develop the wind tunnel model, but additional tailoring is also conducted such that the wind tunnel model achieves a 10% wing tip deflection at the wind tunnel test condition. An aeroelastic scaling procedure and analysis is conducted, and a sub-scale flexible wind tunnel model based on the full-scale's undeformed jig-shape is developed. Optimization of the flexible wind tunnel model's undeflected twist along the span, or pre-twist or wash-out, is then conducted for the design test condition. The resulting wind tunnel model is an aeroelastic model designed for the wind tunnel test condition.

Power of the wingbeat: modelling the effects of flapping wings in vertebrate flight.

PubMed

Heerenbrink, M Klein; Johansson, L C; Hedenström, A

2015-05-08

Animal flight performance has been studied using models developed for man-made aircraft. For an aeroplane with fixed wings, the energetic cost as a function of flight speed can be expressed in terms of weight, wing span, wing area and body area, where more details are included in proportionality coefficients. Flying animals flap their wings to produce thrust. Adopting the fixed wing flight model implicitly incorporates the effects of wing flapping in the coefficients. However, in practice, these effects have been ignored. In this paper, the effects of reciprocating wing motion on the coefficients of the fixed wing aerodynamic power model for forward flight are explicitly formulated in terms of thrust requirement, wingbeat frequency and stroke-plane angle, for optimized wingbeat amplitudes. The expressions are obtained by simulating flights over a large parameter range using an optimal vortex wake method combined with a low-level blade element method. The results imply that previously assumed acceptable values for the induced power factor might be strongly underestimated. The results also show the dependence of profile power on wing kinematics. The expressions introduced in this paper can be used to significantly improve animal flight models.

Power of the wingbeat: modelling the effects of flapping wings in vertebrate flight

PubMed Central

Heerenbrink, M. Klein; Johansson, L. C.; Hedenström, A.

2015-01-01

Animal flight performance has been studied using models developed for man-made aircraft. For an aeroplane with fixed wings, the energetic cost as a function of flight speed can be expressed in terms of weight, wing span, wing area and body area, where more details are included in proportionality coefficients. Flying animals flap their wings to produce thrust. Adopting the fixed wing flight model implicitly incorporates the effects of wing flapping in the coefficients. However, in practice, these effects have been ignored. In this paper, the effects of reciprocating wing motion on the coefficients of the fixed wing aerodynamic power model for forward flight are explicitly formulated in terms of thrust requirement, wingbeat frequency and stroke-plane angle, for optimized wingbeat amplitudes. The expressions are obtained by simulating flights over a large parameter range using an optimal vortex wake method combined with a low-level blade element method. The results imply that previously assumed acceptable values for the induced power factor might be strongly underestimated. The results also show the dependence of profile power on wing kinematics. The expressions introduced in this paper can be used to significantly improve animal flight models. PMID:27547098

Physical properties of the benchmark models program supercritical wing

NASA Technical Reports Server (NTRS)

Dansberry, Bryan E.; Durham, Michael H.; Bennett, Robert M.; Turnock, David L.; Silva, Walter A.; Rivera, Jose A., Jr.

1993-01-01

The goal of the Benchmark Models Program is to provide data useful in the development and evaluation of aeroelastic computational fluid dynamics (CFD) codes. To that end, a series of three similar wing models are being flutter tested in the Langley Transonic Dynamics Tunnel. These models are designed to simultaneously acquire model response data and unsteady surface pressure data during wing flutter conditions. The supercritical wing is the second model of this series. It is a rigid semispan model with a rectangular planform and a NASA SC(2)-0414 supercritical airfoil shape. The supercritical wing model was flutter tested on a flexible mount, called the Pitch and Plunge Apparatus, that provides a well-defined, two-degree-of-freedom dynamic system. The supercritical wing model and associated flutter test apparatus is described and experimentally determined wind-off structural dynamic characteristics of the combined rigid model and flexible mount system are included.

The effect of wing flexibility on sound generation of flapping wings.

PubMed

Geng, Biao; Xue, Qian; Zheng, Xudong; Liu, Geng; Ren, Yan; Dong, Haibo

2017-12-13

In this study, the unsteady flow and acoustic characteristics of a three-dimensional (3D) flapping wing model of a Tibicen linnei cicada in forward-flight are numerically investigated. A single cicada wing is modelled as a membrane with a prescribed motion reconstructed from high-speed videos of a live insect. The numerical solution takes a hydrodynamic/acoustic splitting approach: the flow field is solved with an incompressible Navier-Stokes flow solver based on an immersed boundary method, and the acoustic field is solved with linearized perturbed compressible equations. The 3D simulation allows for the examination of both the directivity and frequency compositions of the flapping wing sound in a full space. Along with the flexible wing model, a rigid wing model that is extracted from real motion is also simulated to investigate the effects of wing flexibility. The simulation results show that the flapping sound is directional; the dominant frequency varies around the wing. The first and second frequency harmonics show different radiation patterns in the rigid and flexible wing cases, which are demonstrated to be highly associated with wing kinematics and loadings. Furthermore, the rotation and deformation in the flexible wing is found to help lower the sound strength in all directions.

Store Separation Simulation of the Penguin Missile from Helicopters

DTIC Science & Technology

2006-05-01

Fin Sections – Parent Aircraft Aerodynamic Modeling • Fuselage • Wing and Pylon – Flight Simulation Features • Eqns. Of Motion • Ejectors , Thrust ...model – Lanyard model – Models for ejectors , thrust , mass, etc… – Helicopter rotor wake model – Penguin wing deployment dynamics – Penguin wing roll...umbilical, wing roll tabs, time dependent thrust and mass properties, and the incorporation of a realistic autopilot. The modeling of the unique

KSC-2014-3324

NASA Image and Video Library

2014-07-23

VANDENBERG AIR FORCE BASE, Calif. – The first stage of the United Launch Alliance Delta II rocket for NASA's Soil Moisture Active Passive mission, or SMAP, is transported from the Building 836 hangar to the Horizontal Processing Facility at Space Launch Complex 2 on Vandenberg Air Force Base in California. SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch is scheduled for November 2014. To learn more about SMAP, visit http://smap.jpl.nasa.gov. Photo credit: NASA/U.S. Air Force 30th Space Wing

High Speed Civil Transport Design Using Collaborative Optimization and Approximate Models

NASA Technical Reports Server (NTRS)

Manning, Valerie Michelle

1999-01-01

The design of supersonic aircraft requires complex analysis in multiple disciplines, posing, a challenge for optimization methods. In this thesis, collaborative optimization, a design architecture developed to solve large-scale multidisciplinary design problems, is applied to the design of supersonic transport concepts. Collaborative optimization takes advantage of natural disciplinary segmentation to facilitate parallel execution of design tasks. Discipline-specific design optimization proceeds while a coordinating mechanism ensures progress toward an optimum and compatibility between disciplinary designs. Two concepts for supersonic aircraft are investigated: a conventional delta-wing design and a natural laminar flow concept that achieves improved performance by exploiting properties of supersonic flow to delay boundary layer transition. The work involves the development of aerodynamics and structural analyses, and integration within a collaborative optimization framework. It represents the most extensive application of the method to date.

KSC-2014-3325

NASA Image and Video Library

2014-07-23

VANDENBERG AIR FORCE BASE, Calif. – The first stage of the United Launch Alliance Delta II rocket for NASA's Soil Moisture Active Passive mission, or SMAP, makes its way along the roadways on Vandenberg Air Force Base in California from the Building 836 hangar to the Horizontal Processing Facility at Space Launch Complex 2. SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch is scheduled for November 2014. To learn more about SMAP, visit http://smap.jpl.nasa.gov. Photo credit: NASA/U.S. Air Force 30th Space Wing

KSC-2011-2195

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --At Vandenberg Air Force Base in California, a crane raises one of three United Launch Alliance Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2193

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --As the sun rises over Vandenberg Air Force Base in California, a crane raises one of three United Launch Alliance Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2190

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --As the sun rises over Vandenberg Air Force Base in California, United Launch Alliance technicians prepare to raise one of three Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2192

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --As the sun rises over Vandenberg Air Force Base in California, a crane raises one of three United Launch Alliance Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2189

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --As the sun rises over Vandenberg Air Force Base in California, United Launch Alliance technicians prepare to raise one of three Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2198

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --At Vandenberg Air Force Base in California, United Launch Alliance technicians finish installing one of three Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2014-1984

NASA Image and Video Library

2014-03-20

VANDENBERG AIR FORCE BASE, Calif. – The Delta first-stage booster for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, winds its way along the roads at Vandenberg Air Force Base in California on its move from the Building 836 hangar to the Horizontal Processing Facility at Space Launch Complex 2. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket on July 1, 2014. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/D. Liberotti, 30th Space Wing, VAFB

KSC-2014-1987

NASA Image and Video Library

2014-03-20

VANDENBERG AIR FORCE BASE, Calif. – The Delta first-stage booster for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, is escorted along the roads at Vandenberg Air Force Base in California on its move from the Building 836 hangar to the Horizontal Processing Facility at Space Launch Complex 2. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket on July 1, 2014. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/D. Liberotti, 30th Space Wing, VAFB

KSC-2014-1991

NASA Image and Video Library

2014-03-20

VANDENBERG AIR FORCE BASE, Calif. – The Delta first-stage booster for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, rolls into position through the open door of the Horizontal Processing Facility at Space Launch Complex 2 on Vandenberg Air Force Base in California. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket on July 1, 2014. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/D. Liberotti, 30th Space Wing, VAFB

Density effects on the electronic contribution to hydrogen Lyman alpha Stark profiles

NASA Astrophysics Data System (ADS)

Motapon, O.

1998-01-01

The quantum unified theory of Stark broadening (Tran Minh et al. 1975, Feautrier et al. 1976) is used to study the density effects on the electronic contribution to the hydrogen Lyman alpha lineshape. The contribution of the first angular momenta to the total profile is obtained by an extrapolation method, and the results agree with other approaches. The comparison made with Vidal et al. (1973) shows a good agreement; and the electronic profile is found to be linear in density for | Delta lambda right | greater than 8 Angstroms for densities below 10(17) cm(-3) , while the density dependence becomes more complex for | Delta lambda right | less than 8 Angstroms. The wing profiles are calculated at various temperatures scaling from 2500 to 40000K and a polynomial fit of these profiles is given.

Wind tunnel pressure study and Euler code validation of a missile configuration with 77 deg swept delta wings at supersonic speeds. M.S. Thesis - George Washington Univ.

NASA Technical Reports Server (NTRS)

Fulton, Patsy S.

1988-01-01

A wind-tunnel pressure study was conducted on an axisymmetric missile configuration in the Unitary Plan Wind Tunnel at NASA Langley Research Center. The Mach numbers ranged from 1.70 to 2.86 and the angles of attack ranged from minus 4 degrees to plus 24 degrees. The computational accuracy for limited conditions of a space-marching Euler code was assessed.

Surface pressure and inviscid flow field properties of the North American Rockwell delta-wing orbiter for nominal Mach number of 8, volume 2

NASA Technical Reports Server (NTRS)

Matthews, R. K.; Martindale, W. R.; Warmbrod, J. D.

1972-01-01

The results are presented of a wind tunnel test program to determine surface pressures and flow field properties on the space shuttle orbiter configuration. The tests were conducted in September 1971. Data were obtained at a nominal Mach number of 8 and a free stream unit Reynolds number of 3.7 million per foot. Angle of attack was varied from 10 to 50 deg in 10-deg increments.

NACA Aircraft in hangar 1953 - clockwise from front center: YF-84A, D-558-1, D-558-2, B-47, X-1 ship

NASA Technical Reports Server (NTRS)

1953-01-01

In the center foreground of this 1953 hanger photo is the YF-84A (NACA 134/Air Force 45-59490) used for vortex generator research. It arrived on November 28, 1949, and departed on April 21, 1954. Beside it is the third D-558-1 aircraft (NACA 142/Navy 37972). This aircraft was used for a total of 78 transonic research flights from April 1949 to June 1954. It replaced the second D-558-1, lost in the crash which killed Howard Lilly. Just visible on the left edge is the nose of the first D-558-2 (NACA 143/Navy 37973). Douglas turned the aircraft over to NACA on August 31, 1951, after the contractor had completed its initial test flights. NACA only made a single flight with the aircraft, on September 17, 1956, before the program was cancelled. In the center of the photo is the B-47A (NACA 150/Air Force 49-1900). The B-47 jet bomber, with its thin, swept-back wings, and six podded engines, represented the state of the art in aircraft design in the early 1950s. The aircraft undertook a number of research activities between May 1953 and its 78th and final research flight on November 22, 1957. The tests showed that the aircraft had a buffeting problem at speeds above Mach 0.8. Among the pilots who flew the B-47 were later X-15 pilots Joe Walker, A. Scott Crossfield, John B. McKay, and Neil A. Armstrong. On the right side of the B-47 is NACA's X-1 (Air Force 46-063). The second XS-1 aircraft built, it was fitted with a thicker wing than that on the first aircraft, which had exceeded Mach 1 on October 14, 1947. Flight research by NACA pilots indicated that this thicker wing produced 30 percent more drag at transonic speeds compared to the thinner wing on the first X-1. After a final flight on October 23, 1951, the aircraft was grounded due to the possibility of fatigue failure of the nitrogen spheres used to pressurize the fuel tanks. At the time of this photo, in 1953, the aircraft was in storage. In 1955, the aircraft was extensively modified, becoming the X-1E. In front of the X-1 is the XF-92A (Air Force 46-682). Unlike the X-1 and D-558 aircraft, the XF-92A was not designed as a research aircraft, but as the prototype of a delta-wing fighter. While the effort was unsuccessful, the XF-92A offered the chance to test a delta wing aircraft. A brief series of 25 flights were made using the aircraft in 1953. These showed the aircraft had violent pitch-up tendencies during turns. Despite the problems, the XF-92A contributed to later delta wing aircraft, like the F-102, F-106, and B-58. Behind the B-47, in the back of the hangar, are four other aircraft. From left to right, they are the second X-4 (Air Force 46-677) research aircraft. It was operated by the NACA from May 8, 1950, to March 22, 1954, when it left the High-Speed Flight Research Station for the U.S. Air Force Museum. It was designed to test the use of swept wings but no horizontal stabilizers. This proved to have poor transonic stability. Next to it is the ETF-51D Mustang (NACA 148/Air Force 44-84958) used for low-speed chase missions, as well as support and liaison flights. On the right side of the B-47 is the first D-558-1. Originally given the Navy number 37970, it was flown as part of the Douglas contractor program. When this was completed, the aircraft was turned over to the NACA on April 11, 1949. Although the aircraft was designation 'NACA 140,' it was never flown again. Instead, it was used to provide spare parts to keep the third D-558-1 in operation. In this photo the aircraft is partially disassembled. The final aircraft is the first X-5 (Air Force 50-1838). This was a research aircraft used to test the concept of pivoting wings which could change their sweep angle in flight. The results were mixed; the X-5 had vicious stall behavior due to the poor position of the tail and stabilizers. The mechanism used by the 'variable-sweep wing' was also complex, which limited its usefulness. Despite these problems, the X-5's primary advantage was that it was equivalent to a whole family of research aircraft. It could provide transonic data at sweep angles up to 60 degrees--the same as the delta wing XF-92A. The Dryden Flight Research Center, NASA's premier installation for aeronautical flight research, celebrated its 50th anniversary in 1996. Dryden is the 'Center of Excellence' for atmospheric flight operations. The Center's charter is to research, develop, verify, and transfer advanced aeronautics, space, and related technologies. It is located at Edwards, Calif., on the western edge of the Mojave Desert, 80 miles north of Los Angeles. Dryden's history dates back to the early fall of 1946, when a group of five aeronautical engineers arrived at what is now Edwards from the NACA's Langley Memorial Aeronautical Laboratory, Hampton, Va. Their goal was to prepare for the X-l supersonic research flights in a joint NACA-U.S. Army Air Forces-Bell Aircraft Corp. program. NACA--the National Advisory Committee for Aeronautics--was the predecessor of today's NASA. Since the days of the X-l, the first aircraft to fly faster than the speed of sound, the installation has grown in size and significance and is associated with many important developments in aviation -- supersonic and hypersonic flight, wingless lifting bodies, digital fly-by-wire, supercritical and forward-swept wings, and the space shuttles. Its name has changed many times over the years. From 14 November 1949 to 1 July 1954 it bore the name NACA High-Speed Flight Research Station.

Aerodynamic forces and flow structures of the leading edge vortex on a flapping wing considering ground effect.

PubMed

Van Truong, Tien; Byun, Doyoung; Kim, Min Jun; Yoon, Kwang Joon; Park, Hoon Cheol

2013-09-01

The aim of this work is to provide an insight into the aerodynamic performance of the beetle during takeoff, which has been estimated in previous investigations. We employed a scaled-up electromechanical model flapping wing to measure the aerodynamic forces and the three-dimensional flow structures on the flapping wing. The ground effect on the unsteady forces and flow structures were also characterized. The dynamically scaled wing model could replicate the general stroke pattern of the beetle's hind wing kinematics during takeoff flight. Two wing kinematic models have been studied to examine the influences of wing kinematics on unsteady aerodynamic forces. In the first model, the angle of attack is asymmetric and varies during the translational motion, which is the flapping motion of the beetle's hind wing. In the second model, the angle of attack is constant during the translational motion. The instantaneous aerodynamic forces were measured for four strokes during the beetle's takeoff by the force sensor attached at the wing base. Flow visualization provided a general picture of the evolution of the three-dimensional leading edge vortex (LEV) on the beetle hind wing model. The LEV is stable during each stroke, and increases radically from the root to the tip, forming a leading-edge spiral vortex. The force measurement results show that the vertical force generated by the hind wing is large enough to lift the beetle. For the beetle hind wing kinematics, the total vertical force production increases 18.4% and 8.6% for the first and second strokes, respectively, due to the ground effect. However, for the model with a constant angle of attack during translation, the vertical force is reduced during the first stroke. During the third and fourth strokes, the ground effect is negligible for both wing kinematic patterns. This finding suggests that the beetle's flapping mechanism induces a ground effect that can efficiently lift its body from the ground during takeoff.

Aerodynamic effects of corrugation and deformation in flapping wings of hovering hoverflies.

PubMed

Du, Gang; Sun, Mao

2012-05-07

We investigated the aerodynamic effects of wing deformation and corrugation of a three-dimensional model hoverfly wing at a hovering condition by solving the Navier-Stokes equations on a dynamically deforming grid. Various corrugated wing models were tested. Insight into whether or not there existed significant aerodynamic coupling between wing deformation (camber and twist) and wing corrugation was obtained by comparing aerodynamic forces of four cases: a smooth-plate wing in flapping motion without deformation (i.e. a rigid flat-plate wing in flapping motion); a smooth-plate wing in flapping motion with deformation; a corrugated wing in flapping motion without deformation (i.e. a rigid corrugated wing in flapping motion); a corrugated wing in flapping motion with deformation. There was little aerodynamic coupling between wing deformation and corrugation: the aerodynamic effect of wing deformation and corrugation acting together was approximately a superposition of those of deformation and corrugation acting separately. When acting alone, the effect of wing deformation was to increase the lift by 9.7% and decrease the torque (or aerodynamic power) by 5.2%, and that of wing corrugation was to decrease the lift by 6.5% and increase the torque by 2.2%. But when acting together, the wing deformation and corrugation only increased the lift by ~3% and decreased the torque by ~3%. That is, the combined aerodynamic effect of deformation and corrugation is rather small. Thus, wing corrugation is mainly for structural, not aerodynamic, purpose, and in computing or measuring the aerodynamic forces, using a rigid flat-plate wing to model the corrugated deforming wing at hovering condition can be a good approximation. Copyright © 2012 Elsevier Ltd. All rights reserved.

Experimental study of a generic high-speed civil transport: Tabulated data

NASA Technical Reports Server (NTRS)

Belton, Pamela S.; Campbell, Richard L.

1992-01-01

An experimental study of a generic high-speed civil transport was conducted in LaRC's 8-Foot Transonic Pressure Tunnel. The data base was obtained for the purpose of assessing the accuracy of various levels of computational analysis. Two models differing only in wing tip geometry were tested with and without flow-through nacelles. The baseline model has a curved or crescent wing tip shape while the second model has a more conventional straight wing tip shape. The study was conducted at Mach numbers from 0.30-1.19. Force data were obtained on both the straight and curved wing tip models. Only the curved wing tip model was instrumented for measuring pressures. Longitudinal and lateral-directional aerodynamic data are presented without analysis in tabulated form. Pressure coefficients for the curved wing tip model are also presented in tabulated form.

Determination of longitudinal aerodynamic derivatives using flight data from an icing research aircraft

NASA Technical Reports Server (NTRS)

Ranaudo, R. J.; Batterson, J. G.; Reehorst, A. L.; Bond, T. H.; Omara, T. M.

1989-01-01

A flight test was performed with the NASA Lewis Research Center's DH-6 icing research aircraft. The purpose was to employ a flight test procedure and data analysis method, to determine the accuracy with which the effects of ice on aircraft stability and control could be measured. For simplicity, flight testing was restricted to the short period longitudinal mode. Two flights were flown in a clean (baseline) configuration, and two flights were flown with simulated horizontal tail ice. Forty-five repeat doublet maneuvers were performed in each of four test configurations, at a given trim speed, to determine the ensemble variation of the estimated stability and control derivatives. Additional maneuvers were also performed in each configuration, to determine the variation in the longitudinal derivative estimates over a wide range of trim speeds. Stability and control derivatives were estimated by a Modified Stepwise Regression (MSR) technique. A measure of the confidence in the derivative estimates was obtained by comparing the standard error for the ensemble of repeat maneuvers, to the average of the estimated standard errors predicted by the MSR program. A multiplicative relationship was determined between the ensemble standard error, and the averaged program standard errors. In addition, a 95 percent confidence interval analysis was performed for the elevator effectiveness estimates, C sub m sub delta e. This analysis identified the speed range where changes in C sub m sub delta e could be attributed to icing effects. The magnitude of icing effects on the derivative estimates were strongly dependent on flight speed and aircraft wing flap configuration. With wing flaps up, the estimated derivatives were degraded most at lower speeds corresponding to that configuration. With wing flaps extended to 10 degrees, the estimated derivatives were degraded most at the higher corresponding speeds. The effects of icing on the changes in longitudinal stability and control derivatives were adequately determined by the flight test procedure and the MSR analysis method discussed herein.

Detailed pressure distribution measurements obtained on several configurations of an aspect-ratio-7 variable twist wing

NASA Technical Reports Server (NTRS)

Holbrook, G. T.; Dunham, D. M.

1985-01-01

Detailed pressure distribution measurements were made for 11 twist configurations of a unique, multisegmented wing model having an aspect ratio of 7 and a taper ratio of 1. These configurations encompassed span loads ranging from that of an untwisted wing to simple flapped wings both with and without upper-surface spoilers attached. For each of the wing twist configurations, electronic scanning pressure transducers were used to obtain 580 surface pressure measurements over the wing in about 0.1 sec. Integrated pressure distribution measurements compared favorably with force-balance measurements of lift on the model when the model centerbody lift was included. Complete plots and tabulations of the pressure distribution data for each wing twist configuration are provided.

A Method for Computing the Core Flow in Three-Dimensional Leading-Edge Vortices. Ph.D. Thesis - North Carolina State Univ.

NASA Technical Reports Server (NTRS)

Luckring, J. M.

1985-01-01

A theory is presented for calculating the flow in the core of a separation-induced leading-edge vortex. The method is based on matching inner and outer representations of the vortex. The inner model of the vortex is based on the quasicylindrical Navier-Stokes equations; the flow is assumed to be steady, axially symmetric, and incompressible and in addition, gradients in the radial direction are assumed to be much larger then gradients in the axial direction. The outer model is based on the three-dimensional free-vortex-sheet theory, a higher-order panel method which solves the Prandtl-Glauert equation including nonlinear boundary conditions pertinent to the concentrated vorticity representation of the leading edge vortex. The resultant flow is evaluated a posteriori for evidence of incipient vortex breakdown and the critical helix angle concept, in conjunction with an adverse longitudinal pressure gradient, is found to correlate well with the occurrence of vortex breakdown at the trailing edge of delta, arrow, and diamond wings.

Cart3D Simulations for the First AIAA Sonic Boom Prediction Workshop

NASA Technical Reports Server (NTRS)

Aftosmis, Michael J.; Nemec, Marian

2014-01-01

Simulation results for the First AIAA Sonic Boom Prediction Workshop (LBW1) are presented using an inviscid, embedded-boundary Cartesian mesh method. The method employs adjoint-based error estimation and adaptive meshing to automatically determine resolution requirements of the computational domain. Results are presented for both mandatory and optional test cases. These include an axisymmetric body of revolution, a 69deg delta wing model and a complete model of the Lockheed N+2 supersonic tri-jet with V-tail and flow through nacelles. In addition to formal mesh refinement studies and examination of the adjoint-based error estimates, mesh convergence is assessed by presenting simulation results for meshes at several resolutions which are comparable in size to the unstructured grids distributed by the workshop organizers. Data provided includes both the pressure signals required by the workshop and information on code performance in both memory and processing time. Various enhanced techniques offering improved simulation efficiency will be demonstrated and discussed.

MODEL TESTS AND 3D ELASTIC FINITE ELEMENT ANALYSIS FOR STEEL PIPE PILES WITH WINGS IN STALLED IN SOIL CEMENT COLUMN

NASA Astrophysics Data System (ADS)

Tamai, Toshiyuki; Teramoto, Shuntarou; Kimura, Makoto

Steel pipe piles with wings installed in soil cement column is a composite foundation of pile consisting of soil improvement with cement and steel pipe with wings. This type of pile shows higher vertical bearing capacity when compared to steel pipe piles that are installed without soil cement. It is thought the wings contribute to higher bearing capacity of this type of piles. The wings are also thought to play the role of structural unification of pile foundations and load transfer. In this study, model test and 3D elastic finite element analysis was carried out in order to elucidate the effect of wings on the structural unification of pile foundation and the load transfer mechanism. Firstly, the model test was carried out in order to grasp the influence of pile with and without wings, the shape of wings of the pile and the unconfined compression strength of the soil cement on the structural unification of the pile foundation. The numerical analysis of the model test was then carried out on the intermediate part of the pile foundation with wings and mathematical model developed. Finally load tran sfer mechanism was checked for the entire length of the pile through this mathematical model and the load sharing ratio of the wings and stress distribution occurring in the soil cement clarified. In addition, the effect of the wing interval on the structural unification of the pile foundation and load transfer was also checked and clarified.

Aerodynamic performance and particle image velocimetery of piezo actuated biomimetic manduca sexta engineered wings towards the design and application of a flapping wing flight vehicle

NASA Astrophysics Data System (ADS)

DeLuca, Anthony M.

Considerable research and investigation has been conducted on the aerodynamic performance, and the predominate flow physics of the Manduca Sexta size of biomimetically designed and fabricated wings as part of the AFIT FWMAV design project. Despite a burgeoning interest and research into the diverse field of flapping wing flight and biomimicry, the aerodynamics of flapping wing flight remains a nebulous field of science with considerable variance into the theoretical abstractions surrounding aerodynamic mechanisms responsible for aerial performance. Traditional FWMAV flight models assume a form of a quasi-steady approximation of wing aerodynamics based on an infinite wing blade element model (BEM). An accurate estimation of the lift, drag, and side force coefficients is a critical component of autonomous stability and control models. This research focused on two separate experimental avenues into the aerodynamics of AFIT's engineered hawkmoth wings|forces and flow visualization. 1. Six degree of freedom force balance testing, and high speed video analysis was conducted on 30°, 45°, and 60° angle stop wings. A novel, non-intrusive optical tracking algorithm was developed utilizing a combination of a Gaussian Mixture Model (GMM) and ComputerVision (OpenCV) tools to track the wing in motion from multiple cameras. A complete mapping of the wing's kinematic angles as a function of driving amplitude was performed. The stroke angle, elevation angle, and angle of attack were tabulated for all three wings at driving amplitudes ranging from A=0.3 to A=0.6. The wing kinematics together with the force balance data was used to develop several aerodynamic force coefficient models. A combined translational and rotational aerodynamic model predicted lift forces within 10%, and vertical forces within 6%. The total power consumption was calculated for each of the three wings, and a Figure of Merit was calculated for each wing as a general expression of the overall efficiency of the wing. Th 60° angle stop wing achieved the largest total stroke angle and generated the most lift for the lowest power consumption of the wings tested. 2. Phase averaged stereo Particle Image Velocimetry (PIV) data was collected at eight phases through the flap cycle on the 30°, 45°, and 60° angle stop wings. Wings were mounted transverse and parallel to the interrogating laser sheet, and planar velocity intersections at the wing mid-span, one chord below the wing, were compared to one another to verify data fidelity. A Rankine-Froude actuator disk model was adapted to calculate the approximate vertical thrust generated from the total momentum flux through the flapping semi-disk using the velocity field measurements. Three component stereo u, v, and w-velocity contour measurements confirmed the presence of extensive vortical structures in the vicinity of the wing. The leading edge vortex was successfully tracked through the stroke cycle appearing at approximately 25% span, increasing in circulatory strength and translational velocity down the span toward the tip, and dissipating just after 75% span. Thrust calculations showed the vertically mounted wing more accurately represented the vertical forces when compared to its corresponding force balance measurement than the horizontally mounted wing. The mid-span showed the highest vertical velocity profile below the wing; and hence, was the location responsible for the majority of lift production along the span.

A Model for Selection of Eyespots on Butterfly Wings.

PubMed

Sekimura, Toshio; Venkataraman, Chandrasekhar; Madzvamuse, Anotida

2015-01-01

The development of eyespots on the wing surface of butterflies of the family Nympalidae is one of the most studied examples of biological pattern formation.However, little is known about the mechanism that determines the number and precise locations of eyespots on the wing. Eyespots develop around signaling centers, called foci, that are located equidistant from wing veins along the midline of a wing cell (an area bounded by veins). A fundamental question that remains unsolved is, why a certain wing cell develops an eyespot, while other wing cells do not. We illustrate that the key to understanding focus point selection may be in the venation system of the wing disc. Our main hypothesis is that changes in morphogen concentration along the proximal boundary veins of wing cells govern focus point selection. Based on previous studies, we focus on a spatially two-dimensional reaction-diffusion system model posed in the interior of each wing cell that describes the formation of focus points. Using finite element based numerical simulations, we demonstrate that variation in the proximal boundary condition is sufficient to robustly select whether an eyespot focus point forms in otherwise identical wing cells. We also illustrate that this behavior is robust to small perturbations in the parameters and geometry and moderate levels of noise. Hence, we suggest that an anterior-posterior pattern of morphogen concentration along the proximal vein may be the main determinant of the distribution of focus points on the wing surface. In order to complete our model, we propose a two stage reaction-diffusion system model, in which an one-dimensional surface reaction-diffusion system, posed on the proximal vein, generates the morphogen concentrations that act as non-homogeneous Dirichlet (i.e., fixed) boundary conditions for the two-dimensional reaction-diffusion model posed in the wing cells. The two-stage model appears capable of generating focus point distributions observed in nature. We therefore conclude that changes in the proximal boundary conditions are sufficient to explain the empirically observed distribution of eyespot focus points on the entire wing surface. The model predicts, subject to experimental verification, that the source strength of the activator at the proximal boundary should be lower in wing cells in which focus points form than in those that lack focus points. The model suggests that the number and locations of eyespot foci on the wing disc could be largely controlled by two kinds of gradients along two different directions, that is, the first one is the gradient in spatially varying parameters such as the reaction rate along the anterior-posterior direction on the proximal boundary of the wing cells, and the second one is the gradient in source values of the activator along the veins in the proximal-distal direction of the wing cell.

Flow visualization study of close-coupled canard wing and strake wing configuration

NASA Technical Reports Server (NTRS)

Miner, D. D.; Gloss, B. B.

1975-01-01

The Langley 1/8-scale V/STOL model tunnel was used to qualitatively determine the flow fields associated with semi-span close coupled canard wing and strake wing models. Small helium filled bubbles were injected upstream of the models to make the flow visible. Photographs were taken over the angle-of-attack ranges of -10 deg to 40 deg.

Aerodynamics on a transport aircraft type wing-body model

NASA Technical Reports Server (NTRS)

Schmitt, V.

1982-01-01

The DFLR-F4 wing-body combination is studied. The 1/38 model is formed by a 9.5 aspect ratio transonic wing and an Airbus A 310 fuselage. The F4 wing geometrical characteristics are described and the main experimental results obtained in the S2MA wind tunnel are discussed. Both wing-fuselage interferences and viscous effects, which are important on the wing due to a high rear loading, are investigated by performing 3D calculations. An attempt is made to find their limitations.

Wing-wake interaction destabilizes hover equilibrium of a flapping insect-scale wing.

PubMed

Bluman, James; Kang, Chang-Kwon

2017-06-15

Wing-wake interaction is a characteristic nonlinear flow feature that can enhance unsteady lift in flapping flight. However, the effects of wing-wake interaction on the flight dynamics of hover are inadequately understood. We use a well-validated 2D Navier-Stokes equation solver and a quasi-steady model to investigate the role of wing-wake interaction on the hover stability of a fruit fly scale flapping flyer. The Navier-Stokes equations capture wing-wake interaction, whereas the quasi-steady models do not. Both aerodynamic models are tightly coupled to a flight dynamic model, which includes the effects of wing mass. The flapping amplitude, stroke plane angle, and flapping offset angle are adjusted in free flight for various wing rotations to achieve hover equilibrium. We present stability results for 152 simulations which consider different kinematics involving the pitch amplitude and pitch axis as well as the duration and timing of pitch rotation. The stability of all studied motions was qualitatively similar, with an unstable oscillatory mode present in each case. Wing-wake interaction has a destabilizing effect on the longitudinal stability, which cannot be predicted by a quasi-steady model. Wing-wake interaction increases the tendency of the flapping flyer to pitch up in the presence of a horizontal velocity perturbation, which further destabilizes the unstable oscillatory mode of hovering flight dynamics.

Structural modeling and optimization of a joined-wing configuration of a High-Altitude Long-Endurance (HALE) aircraft

NASA Astrophysics Data System (ADS)

Kaloyanova, Valentina B.

Recent research trends have indicated an interest in High-Altitude, Long-Endurance (HALE) aircraft as a low-cost alternative to certain space missions, such as telecommunication relay, environmental sensing and military reconnaissance. HALE missions require a light vehicle flying at low speed in the stratosphere at altitudes of 60,000-80,000 ft, with a continuous loiter time of up to several days. To provide high lift and low drag at these high altitudes, where the air density is low, the wing area should be increased, i.e., high-aspect-ratio wings are necessary. Due to its large span and lightweight, the wing structure is very flexible. To reduce the structural deformation, and increase the total lift in a long-spanned wing, a sensorcraft model with a joined-wing configuration, proposed by AFRL, is employed. The joined-wing encompasses a forward wing, which is swept back with a positive dihedral angle, and connected with an aft wing, which is swept forward. The joined-wing design combines structural strength, high aerodynamic performance and efficiency. As a first step to study the joined-wing structural behavior an 1-D approximation model is developed. The 1-D approximation is a simple structural model created using ANSYS BEAM4 elements to present a possible approach for the aerodynamics-structure coupling. The pressure loads from the aerodynamic analysis are integrated numerically to obtain the resultant aerodynamic forces and moments (spanwise lift and pitching moment distributions, acting at the aerodynamic center). These are applied on the 1-D structural model. A linear static analysis is performed under this equivalent load, and the deformed shape of the 1-D model is used to obtain the deformed shape of the actual 3-D joined wing, i.e. deformed aerodynamic surface grid. To date in the existing studies, only simplified structural models have been examined. In the present work, in addition to the simple 1-D beam model, a semi-monocoque structural model is developed. All stringers, skin panels, ribs and spars are represented by appropriate elements in a finite-element model. Also, the model accounts for the fuel weight and sensorcraft antennae housed within the wings. Linear and nonlinear static analyses under the aerodynamic load are performed. The stress distribution in the wing as well as deformation is explored. Starting with a structural model with uniform mass distribution, a design optimization is performed to achieve a fully stressed design. As the joined-wing structure is prone to buckling, after the design optimization is complete linear and nonlinear bucking analyses are performed to study the global joined-wing structural instability, the load magnitude at which it is expected to occur, and the buckling mode. The buckled shape of the aft wing (which is subjected to compression) is found to resemble that of a fixed-pinned column. The linear buckling analysis overestimates the buckling load. However, even the nonlinear buckling analysis results in a load factor higher than 3, i.e. the wing structure is buckling safe under its current loading conditions. As the region of the joint has a very complicated geometry that has adverse effects in the flow and stress behavior an independent, more finely meshed model (submodel) of the joint region is generated and analyzed. A detailed discussion of the stress distribution obtained in the joint region via the submodeling technique is presented in this study as well. It is found out that compared to its structural response, the joint adverse effects are much more pronounced in its aerodynamic response, so it is suggested for future studies the geometry of the joint to be optimized based on its aerodynamic performance. As this design and analysis study is aimed towards developing a realistic structural representation of the innovative joined-wing configuration, in addition to the "global", or upper-level optimization, a local level design optimization is performed as well. At the lower (local) level detailed models of wing structural panels are used to compute more complex failure modes and to design the details that are not included in the upper (global) level model. Proper coordination between local skin-stringer panel models and the global joined-wing model prevents inconsistency between the upper- (global) and lower- (local) level design models. (Abstract shortened by UMI.)

Study on flow over finite wing with respect to F-22 raptor, Supermarine Spitfire, F-7 BG aircraft wing and analyze its stability performance and experimental values

NASA Astrophysics Data System (ADS)

Ali, Md. Nesar; Alam, Mahbubul

2017-06-01

A finite wing is a three-dimensional body, and consequently the flow over the finite wing is three-dimensional; that is, there is a component of flow in the span wise direction. The physical mechanism for generating lift on the wing is the existence of a high pressure on the bottom surface and a low pressure on the top surface. The net imbalance of the pressure distribution creates the lift. As a by-product of this pressure imbalance, the flow near the wing tips tends to curl around the tips, being forced from the high-pressure region just underneath the tips to the low-pressure region on top. This flow around the wing tips is shown in the front view of the wing. As a result, on the top surface of the wing, there is generally a span wise component of flow from the tip toward the wing root, causing the streamlines over the top surface to bend toward the root. On the bottom surface of the wing, there is generally a span wise component of flow from the root toward the tip, causing the streamlines over the bottom surface to bend toward the tip. Clearly, the flow over the finite wing is three-dimensional, and therefore we would expect the overall aerodynamic properties of such a wing to differ from those of its airfoil sections. The tendency for the flow to "leak" around the wing tips has another important effect on the aerodynamics of the wing. This flow establishes a circulatory motion that trails downstream of the wing; that is, a trailing vortex is created at each wing tip. The aerodynamics of finite wings is analyzed using the classical lifting line model. This simple model allows a closed-form solution that captures most of the physical effects applicable to finite wings. The model is based on the horseshoe-shaped vortex that introduces the concept of a vortex wake and wing tip vortices. The downwash induced by the wake creates an induced drag that did not exist in the two-dimensional analysis. Furthermore, as wingspan is reduced, the wing lift slope decreases, and the induced drag increases, reducing overall efficiency. To complement the high aspect ratio wing case, a slender wing model is formulated so that the lift and drag can be estimated for this limiting case as well. We analyze the stability performance of F-22 raptor, Supermarine Spitfire, F-7 BG Aircraft wing by using experimental method and simulation software. The experimental method includes fabrication of F-22 raptor, Supermarine Spitfire, F-7 BG Aircraft wing which making material is Gamahr wood. Testing this model wing in wind tunnel test and after getting expected data we also compared this value with analyzing software data for furthermore experiment.

Time-varying wing-twist improves aerodynamic efficiency of forward flight in butterflies.

PubMed

Zheng, Lingxiao; Hedrick, Tyson L; Mittal, Rajat

2013-01-01

Insect wings can undergo significant chordwise (camber) as well as spanwise (twist) deformation during flapping flight but the effect of these deformations is not well understood. The shape and size of butterfly wings leads to particularly large wing deformations, making them an ideal test case for investigation of these effects. Here we use computational models derived from experiments on free-flying butterflies to understand the effect of time-varying twist and camber on the aerodynamic performance of these insects. High-speed videogrammetry is used to capture the wing kinematics, including deformation, of a Painted Lady butterfly (Vanessa cardui) in untethered, forward flight. These experimental results are then analyzed computationally using a high-fidelity, three-dimensional, unsteady Navier-Stokes flow solver. For comparison to this case, a set of non-deforming, flat-plate wing (FPW) models of wing motion are synthesized and subjected to the same analysis along with a wing model that matches the time-varying wing-twist observed for the butterfly, but has no deformation in camber. The simulations show that the observed butterfly wing (OBW) outperforms all the flat-plate wings in terms of usable force production as well as the ratio of lift to power by at least 29% and 46%, respectively. This increase in efficiency of lift production is at least three-fold greater than reported for other insects. Interestingly, we also find that the twist-only-wing (TOW) model recovers much of the performance of the OBW, demonstrating that wing-twist, and not camber is key to forward flight in these insects. The implications of this on the design of flapping wing micro-aerial vehicles are discussed.

Time-Varying Wing-Twist Improves Aerodynamic Efficiency of Forward Flight in Butterflies

PubMed Central

Zheng, Lingxiao; Hedrick, Tyson L.; Mittal, Rajat

2013-01-01

Insect wings can undergo significant chordwise (camber) as well as spanwise (twist) deformation during flapping flight but the effect of these deformations is not well understood. The shape and size of butterfly wings leads to particularly large wing deformations, making them an ideal test case for investigation of these effects. Here we use computational models derived from experiments on free-flying butterflies to understand the effect of time-varying twist and camber on the aerodynamic performance of these insects. High-speed videogrammetry is used to capture the wing kinematics, including deformation, of a Painted Lady butterfly (Vanessa cardui) in untethered, forward flight. These experimental results are then analyzed computationally using a high-fidelity, three-dimensional, unsteady Navier-Stokes flow solver. For comparison to this case, a set of non-deforming, flat-plate wing (FPW) models of wing motion are synthesized and subjected to the same analysis along with a wing model that matches the time-varying wing-twist observed for the butterfly, but has no deformation in camber. The simulations show that the observed butterfly wing (OBW) outperforms all the flat-plate wings in terms of usable force production as well as the ratio of lift to power by at least 29% and 46%, respectively. This increase in efficiency of lift production is at least three-fold greater than reported for other insects. Interestingly, we also find that the twist-only-wing (TOW) model recovers much of the performance of the OBW, demonstrating that wing-twist, and not camber is key to forward flight in these insects. The implications of this on the design of flapping wing micro-aerial vehicles are discussed. PMID:23341923

Effects of spanwise flexibility on the performance of flapping flyers in forward flight.

PubMed

Kodali, Deepa; Medina, Cory; Kang, Chang-Kwon; Aono, Hikaru

2017-11-01

Flying animals possess flexible wings that deform during flight. The chordwise flexibility alters the wing shape, affecting the effective angle of attack and hence the surrounding aerodynamics. However, the effects of spanwise flexibility on the locomotion are inadequately understood. Here, we present a two-way coupled aeroelastic model of a plunging spanwise flexible wing. The aerodynamics is modelled with a two-dimensional, unsteady, incompressible potential flow model, evaluated at each spanwise location of the wing. The two-way coupling is realized by considering the transverse displacement as the effective plunge under the dynamic balance of wing inertia, elastic restoring force and aerodynamic force. The thrust is a result of the competition between the enhancement due to wing deformation and induced drag. The results for a purely plunging spanwise flexible wing agree well with experimental and high-fidelity numerical results from the literature. Our analysis suggests that the wing aspect ratio of the abstracted passerine and goose models corresponds to the optimal aeroelastic response, generating the highest thrust while minimizing the power required to flap the wings. At these optimal aspect ratios, the flapping frequency is near the first spanwise natural frequency of the wing, suggesting that these birds may benefit from the resonance to generate thrust. © 2017 The Author(s).

Flutter tests (IS4) of the 0.0125-scale shuttle reflection plane model 30-OTS in the Langley Research Center 26-inch transonic blowdown tunnel test no. 547

NASA Technical Reports Server (NTRS)

Kotch, M. A.

1974-01-01

A series of slab wing flutter models with rigid orbiter fuselage, external tank, and SRB models of the space shuttle were tested, in a reflection plane arrangement, in the NASA Langley Research Center's 26-inch Transonic Blowdown Tunnel. Model flutter boundaries were obtained for both a wing-alone configuration and a wing-with-orbiter, tank and SRB configuration. Additional test points were taken of the wing-with-orbiter configuration, as a correlation with the wing-alone condition. A description of the wind tunnel models and test procedures utilized in the experiment are provided.

Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

NASA Technical Reports Server (NTRS)

Allen, Michael J.; Lizotte, Andrew M.; Dibley, Ryan P.; Clarke, Robert

2005-01-01

The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.

Details of insect wing design and deformation enhance aerodynamic function and flight efficiency.

PubMed

Young, John; Walker, Simon M; Bomphrey, Richard J; Taylor, Graham K; Thomas, Adrian L R

2009-09-18

Insect wings are complex structures that deform dramatically in flight. We analyzed the aerodynamic consequences of wing deformation in locusts using a three-dimensional computational fluid dynamics simulation based on detailed wing kinematics. We validated the simulation against smoke visualizations and digital particle image velocimetry on real locusts. We then used the validated model to explore the effects of wing topography and deformation, first by removing camber while keeping the same time-varying twist distribution, and second by removing camber and spanwise twist. The full-fidelity model achieved greater power economy than the uncambered model, which performed better than the untwisted model, showing that the details of insect wing topography and deformation are important aerodynamically. Such details are likely to be important in engineering applications of flapping flight.

KSC-2011-2196

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --At Vandenberg Air Force Base in California, a crane raises one of three United Launch Alliance Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). A second motor was installed earlier in the morning. Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2194

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --At Vandenberg Air Force Base in California, one of three United Space Alliance Delta II solid rocket motors is atop a tug for the move from the solid motor facility to Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2197

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --At Vandenberg Air Force Base in California, a crane raises one of three United Launch Alliance Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). A second motor was installed earlier in the morning. Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2199

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --At Vandenberg Air Force Base in California, United Launch Alliance technicians finish installing one of three Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). A second motor was installed earlier in the morning. Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2191

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --As the sun rises over Vandenberg Air Force Base in California, a crane begins to raise one of three United Launch Alliance Delta II solid rocket motors on the pad at Space Launch Complex-2 West (SLC-2W). Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2011-2188

NASA Image and Video Library

2011-03-09

VANDENBERG AIR FORCE BASE, Calif. --Before the sun rises over Vandenberg Air Force Base in California, United Launch Alliance technicians prepare to move one of three Delta II solid rocket motors from the solid motor facility to Space Launch Complex-2 West (SLC-2W) atop a tug. ULA technician Eric Chambless is in the tug's driver seat. Scheduled to launch in June, the Delta II rocket will carry NASA's Aquarius satellite into low Earth orbit. Aquarius' mission will be to provide monthly maps of global changes in sea surface salinity. By measuring ocean salinity from space, Aquarius will provide new insights into how the massive natural exchange of freshwater between the ocean, atmosphere and sea ice influences ocean circulation, weather and climate. Also going up with the satellite are optical and thermal cameras, a microwave radiometer and the SAC-D spacecraft, which were developed with the help of institutions in Italy, France, Canada and Argentina. Photo credit: VAFB/30th Space Wing

KSC-2014-1988

NASA Image and Video Library

2014-03-20

VANDENBERG AIR FORCE BASE, Calif. – The Delta first-stage booster for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, negotiates an overpass behind its escort vehicles as it makes its way through the roads on Vandenberg Air Force Base in California on its transfer from the Building 836 hangar to the Horizontal Processing Facility at Space Launch Complex 2. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket on July 1, 2014. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/D. Liberotti, 30th Space Wing, VAFB

OCO-2 - Delta II 2nd Stage Transport

NASA Image and Video Library

2014-02-24

VANDENBERG AIR FORCE BASE, Calif. – Workers push the second stage for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, toward the Horizontal Processing Facility at Space Launch Complex 2 on Vandenberg Air Force Base in California. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket in July. The rocket's second stage will insert OCO-2 into a polar Earth orbit. OCO-2 will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/Arron Tauman, 30th Space Wing, VAFB

KSC-2014-1985

NASA Image and Video Library

2014-03-20

VANDENBERG AIR FORCE BASE, Calif. – The Delta first-stage booster for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, passes a static display of a U.S. Air Force Minuteman III intercontinental ballistic missile, at left, on its move from the Building 836 hangar to the Horizontal Processing Facility at Space Launch Complex 2 on Vandenberg Air Force Base in California. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket on July 1, 2014. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/D. Liberotti, 30th Space Wing, VAFB

OCO-2 - Delta II 2nd Stage Offload

NASA Image and Video Library

2014-02-24

VANDENBERG AIR FORCE BASE, Calif. – As the cover of the transportation trailer is lifted in the high bay of the Building 836 hangar on Vandenberg Air Force Base in California, the second stage for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, comes into view. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket from Space Launch Complex 2 in July. The rocket's second stage will insert OCO-2 into a polar Earth orbit. OCO-2 will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/Arron Tauman, 30th Space Wing, VAFB

OCO-2 Fairings being hoisted into MST

NASA Image and Video Library

2014-03-24

VANDENBERG AIR FORCE BASE, Calif. – Half of the fairing for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, is lifted up the side of the mobile service tower at Space Launch Complex 2 on Vandenberg Air Force Base in California toward the Delta II launcher's environmental enclosure, or clean room, at the top of the tower. The fairing will protect OCO-2 during launch aboard a United Launch Alliance Delta II rocket from Space Launch Complex 2 in July. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/30th Space Wing, VAFB

Prediction of forces and moments for flight vehicle control effectors. Part 2: An analysis of delta wing aerodynamic control effectiveness in ground effect

NASA Technical Reports Server (NTRS)

Maughmer, Mark D.; Ozoroski, L.; Ozoroski, T.; Straussfogel, D.

1990-01-01

Many types of hypersonic aircraft configurations are currently being studied for feasibility of future development. Since the control of the hypersonic configurations throughout the speed range has a major impact on acceptable designs, it must be considered in the conceptual design stage. Here, an investigation of the aerodynamic control effectiveness of highly swept delta planforms operating in ground effect is presented. A vortex-lattice computer program incorporating a free wake is developed as a tool to calculate aerodynamic stability and control derivatives. Data generated using this program are compared to experimental data and to data from other vortex-lattice programs. Results show that an elevon deflection produces greater increments in C sub L and C sub M in ground effect than the same deflection produces out of ground effect and that the free wake is indeed necessary for good predictions near the ground.

OCO-2 Fairings being hoisted into MST

NASA Image and Video Library

2014-03-24

VANDENBERG AIR FORCE BASE, Calif. – Half of the fairing for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, is attached to a crane for its lift into the Delta II launcher's environmental enclosure, or clean room, at the top of the mobile service tower at Space Launch Complex 2 on Vandenberg Air Force Base in California. The fairing will protect OCO-2 during launch aboard a United Launch Alliance Delta II rocket from Space Launch Complex 2 in July. The observatory will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/30th Space Wing, VAFB

OCO-2 Interstage Offload

NASA Image and Video Library

2014-02-10

VANDENBERG AIR FORCE BASE, Calif. – The interstage adapter, or ISA, for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, arrives at the Building 836 hangar on Vandenberg Air Force Base in California. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket from Space Launch Complex 2 in July. The ISA is the interface between the Delta II first and second stages. The second stage engine fits within the ISA. OCO-2 will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/Jeremy Moore, 30th Space Wing, VAFB

OCO-2 Interstage Offload

NASA Image and Video Library

2014-02-10

VANDENBERG AIR FORCE BASE, Calif. – The high bay of the Building 836 hangar on Vandenberg Air Force Base in California is ready to receive the interstage adapter, or ISA, for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket from Space Launch Complex 2 in July. The ISA is the interface between the Delta II first and second stages. The second stage engine fits within the ISA. OCO-2 will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/Jeremy Moore, 30th Space Wing, VAFB

OCO-2 Interstage Offload

NASA Image and Video Library

2014-02-10

VANDENBERG AIR FORCE BASE, Calif. – Workers attach the interstage adapter, or ISA, for NASA's Orbiting Carbon Observatory-2 mission, or OCO-2, to a lifting device in the high bay of the Building 836 hangar on Vandenberg Air Force Base in California. OCO-2 is scheduled to launch aboard a United Launch Alliance Delta II rocket from Space Launch Complex 2 in July. The ISA is the interface between the Delta II first and second stages. The second stage engine fits within the ISA. OCO-2 will collect precise global measurements of carbon dioxide in the Earth's atmosphere and provide scientists with a better idea of the chemical compound's impacts on climate change. Scientists will analyze this data to improve our understanding of the natural processes and human activities that regulate the abundance and distribution of this important atmospheric gas. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/Jeremy Moore, 30th Space Wing, VAFB

KSC-2014-2993

NASA Image and Video Library

2014-06-16

VANDENBERG AIR FORCE BASE, Calif. – Technicians in the Astrotech Payload Processing Facility on Vandenberg Air Force Base in California make final preparations to transport NASA's Orbiting Carbon Observatory-2, or OCO-2, to Space Launch Complex 2 for enclosure in the Delta II payload fairing. Launch aboard a United Launch Alliance Delta II rocket is scheduled for 5:56 a.m. EDT on July 1. OCO-2 is NASA’s first mission dedicated to studying atmospheric carbon dioxide, the leading human-produced greenhouse gas driving changes in Earth’s climate. OCO-2 will provide a new tool for understanding the human and natural sources of carbon dioxide emissions and the natural "sinks" that absorb carbon dioxide and help control its buildup. The observatory will measure the global geographic distribution of these sources and sinks and study their changes over time. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/30th Space Wing, U.S. Air Force

KSC-2014-2990

NASA Image and Video Library

2014-06-16

VANDENBERG AIR FORCE BASE, Calif. – Final preparations are underway in the Astrotech Payload Processing Facility on Vandenberg Air Force Base in California to transport NASA's Orbiting Carbon Observatory-2, or OCO-2, to Space Launch Complex 2 for encapsulation in the Delta II payload fairing. Launch aboard a United Launch Alliance Delta II rocket is scheduled for 5:56 a.m. EDT on July 1. OCO-2 is NASA’s first mission dedicated to studying atmospheric carbon dioxide, the leading human-produced greenhouse gas driving changes in Earth’s climate. OCO-2 will provide a new tool for understanding the human and natural sources of carbon dioxide emissions and the natural "sinks" that absorb carbon dioxide and help control its buildup. The observatory will measure the global geographic distribution of these sources and sinks and study their changes over time. To learn more about OCO-2, visit http://oco.jpl.nasa.gov. Photo credit: NASA/30th Space Wing, U.S. Air Force

Investigation at Low Speeds of Deflectors and Spoilers as Gust Alleviators on a Model of the Bell X-5 Airplane with 35 Degree Swept Wings and on a High-aspect-ratio 35 Degree Swept-wing-fuselage Model

NASA Technical Reports Server (NTRS)

Croom, Delwin R; Huffman, Jarrett K

1957-01-01

Results of an investigation at low speeds to determine the gust-alleviation capabilities (reduction in lift-curve slope) of spoilers and deflectors on a 35 degree swept-wing model of high aspect ratio and on a 1/4-scale model of the X-5 airplane with 35 degree swept wings indicate that deflector and spoiler-deflector types of controls can be designed to provide considerable gust alleviation for a swept-wing airplane while still maintaining stability and control.

Frequencies and Flutter Speed Estimation for Damaged Aircraft Wing Using Scaled Equivalent Plate Analysis

NASA Technical Reports Server (NTRS)

Krishnamurthy, Thiagarajan

2010-01-01

Equivalent plate analysis is often used to replace the computationally expensive finite element analysis in initial design stages or in conceptual design of aircraft wing structures. The equivalent plate model can also be used to design a wind tunnel model to match the stiffness characteristics of the wing box of a full-scale aircraft wing model while satisfying strength-based requirements An equivalent plate analysis technique is presented to predict the static and dynamic response of an aircraft wing with or without damage. First, a geometric scale factor and a dynamic pressure scale factor are defined to relate the stiffness, load and deformation of the equivalent plate to the aircraft wing. A procedure using an optimization technique is presented to create scaled equivalent plate models from the full scale aircraft wing using geometric and dynamic pressure scale factors. The scaled models are constructed by matching the stiffness of the scaled equivalent plate with the scaled aircraft wing stiffness. It is demonstrated that the scaled equivalent plate model can be used to predict the deformation of the aircraft wing accurately. Once the full equivalent plate geometry is obtained, any other scaled equivalent plate geometry can be obtained using the geometric scale factor. Next, an average frequency scale factor is defined as the average ratio of the frequencies of the aircraft wing to the frequencies of the full-scaled equivalent plate. The average frequency scale factor combined with the geometric scale factor is used to predict the frequency response of the aircraft wing from the scaled equivalent plate analysis. A procedure is outlined to estimate the frequency response and the flutter speed of an aircraft wing from the equivalent plate analysis using the frequency scale factor and geometric scale factor. The equivalent plate analysis is demonstrated using an aircraft wing without damage and another with damage. Both of the problems show that the scaled equivalent plate analysis can be successfully used to predict the frequencies and flutter speed of a typical aircraft wing.

Development of a multicomponent force and moment balance for water tunnel applications, volume 1

NASA Technical Reports Server (NTRS)

Suarez, Carlos J.; Malcolm, Gerald N.; Kramer, Brian R.; Smith, Brooke C.; Ayers, Bert F.

1994-01-01

The principal objective of this research effort was to develop a multicomponent strain gauge balance to measure forces and moments on models tested in flow visualization water tunnels. An internal balance was designed that allows measuring normal and side forces, and pitching, yawing and rolling moments (no axial force). The five-components to applied loads, low interactions between the sections and no hysteresis. Static experiments (which are discussed in this Volume) were conducted in the Eidetics water tunnel with delta wings and a model of the F/A-18. Experiments with the F/A-18 model included a thorough baseline study and investigations of the effect of control surface deflections and of several Forebody Vortex Control (FVC) techniques. Results were compared to wind tunnel data and, in general, the agreement is very satisfactory. The results of the static tests provide confidence that loads can be measured accurately in the water tunnel with a relatively simple multicomponent internal balance. Dynamic experiments were also performed using the balance, and the results are discussed in detail in Volume 2 of this report.

The effect of line damping, magneto-optics and parasitic light on the derivation of sunspot vector magnetic fields

NASA Technical Reports Server (NTRS)

Skumanich, A.; Lites, B. W.

1985-01-01

The least square fitting of Stokes observations of sunspots using a Milne-Eddington-Unno model appears to lead, in many circumstances, to various inconsistencies such as anomalously large doppler widths and, hence, small magnetic fields which are significantly below those inferred solely from the Zeeman splitting in the intensity profile. It is found that the introduction of additional physics into the model such as the inclusion of damping wings and magneto-optic birefrigence significantly improves the fit to Stokes parameters. Model fits excluding the intensity profile, i.e., of both magnitude as well as spectral shape of the polarization parameters alone, suggest that parasitic light in the intensity profile may also be a source of inconsistencies. The consequences of the physical changes on the vector properties of the field derived from the Fe I lambda 6173 line for the 17 November 1975 spot as well as on the thermodynamic state are discussed. A Doppler width delta lambda (D) - 25mA is bound to be consistent with a low spot temperature and microturbulence, and a damping constant of a = 0.2.

Transonic wing DFVLR-F4 as European test model

NASA Technical Reports Server (NTRS)

Redeker, G.; Schmidt, N.

1980-01-01

A transonic wing, the DFVLR-F4 was designed and tested as a model in European transonic wind tunnels and was found to give performance improvements over conventional wings. One reason for the improvement was the reduction of compression shocks in the transonic region as the result of improved wing design.

KSC-2014-3326

NASA Image and Video Library

2014-07-23

VANDENBERG AIR FORCE BASE, Calif. – The first stage of the United Launch Alliance Delta II rocket for NASA's Soil Moisture Active Passive mission, or SMAP, crosses a railroad bridge on its move from the Building 836 hangar to the Horizontal Processing Facility at Space Launch Complex 2 on Vandenberg Air Force Base in California. SMAP will provide global measurements of soil moisture and its freeze/thaw state. These measurements will be used to enhance understanding of processes that link the water, energy and carbon cycles, and to extend the capabilities of weather and climate prediction models. SMAP data also will be used to quantify net carbon flux in boreal landscapes and to develop improved flood prediction and drought monitoring capabilities. Launch is scheduled for November 2014. To learn more about SMAP, visit http://smap.jpl.nasa.gov. Photo credit: NASA/U.S. Air Force 30th Space Wing

Coupled nonlinear aeroelasticity and flight dynamics of fully flexible aircraft

NASA Astrophysics Data System (ADS)

Su, Weihua

This dissertation introduces an approach to effectively model and analyze the coupled nonlinear aeroelasticity and flight dynamics of highly flexible aircraft. A reduced-order, nonlinear, strain-based finite element framework is used, which is capable of assessing the fundamental impact of structural nonlinear effects in preliminary vehicle design and control synthesis. The cross-sectional stiffness and inertia properties of the wings are calculated along the wing span, and then incorporated into the one-dimensional nonlinear beam formulation. Finite-state unsteady subsonic aerodynamics is used to compute airloads along lifting surfaces. Flight dynamic equations are then introduced to complete the aeroelastic/flight dynamic system equations of motion. Instead of merely considering the flexibility of the wings, the current work allows all members of the vehicle to be flexible. Due to their characteristics of being slender structures, the wings, tail, and fuselage of highly flexible aircraft can be modeled as beams undergoing three dimensional displacements and rotations. New kinematic relationships are developed to handle the split beam systems, such that fully flexible vehicles can be effectively modeled within the existing framework. Different aircraft configurations are modeled and studied, including Single-Wing, Joined-Wing, Blended-Wing-Body, and Flying-Wing configurations. The Lagrange Multiplier Method is applied to model the nodal displacement constraints at the joint locations. Based on the proposed models, roll response and stability studies are conducted on fully flexible and rigidized models. The impacts of the flexibility of different vehicle members on flutter with rigid body motion constraints, flutter in free flight condition, and roll maneuver performance are presented. Also, the static stability of the compressive member of the Joined-Wing configuration is studied. A spatially-distributed discrete gust model is incorporated into the time simulation of the framework. Gust responses of the Flying-Wing configuration subject to stall effects are investigated. A bilinear torsional stiffness model is introduced to study the skin wrinkling due to large bending curvature of the Flying-Wing. The numerical studies illustrate the improvements of the existing reduced-order formulation with new capabilities of both structural modeling and coupled aeroelastic and flight dynamic analysis of fully flexible aircraft.

International Pacific Air and Space Technology Conference and Aircraft Symposium, 29th, Gifu, Japan, Oct. 7-11, 1991, Proceedings

DOE Office of Scientific and Technical Information (OSTI.GOV)

Not Available

1991-01-01

Various papers on air and space technology are presented. Individual topics addressed include: media selection analysis: implications for training design, high-speed challenge for rotary wing aircraft, high-speed VSTOL answer to congestion, next generation in computational aerodynamics, acrobatic airship 'Acrostat', ducted fan VTOL for working platform, Arianespace launch of Lightsats, small particle acceleration by minirailgun, free-wake analyses of a hovering rotor using panel method, update of the X-29 high-angle-of-attack program, economic approach to accurate wing design, flow field around thick delta wing with rounded leading edge, aerostructural integrated design of forward-swept wing, static characteristics of a two-phase fluid drop system, simplfied-modelmore » approach to group combustion of fuel spray, avionics flight systems for the 21st century. Also discussed are: Aircraft Command in Emergency Situations, spectrogram diagnosis of aircraft disasters, shock interaction induced by two hemisphere-cylinders, impact response of composite UHB propeller blades, high-altitude lighter-than-air powered platform, integrated wiring system, auxiliary power units for current and future aircraft, Space Shuttle Orbiter Auxiliary Power Unit status, numerical analysis of RCS jet in hypersonic flights, energy requirements for the space frontier, electrical system options for space exploration, aerospace plane hydrogen scramjet boosting, manual control of vehicles with time-varying dynamics, design of strongly stabilizing controller, development of the Liquid Apogee Propulsion System for ETS-VI.« less

Subsonic Aerodynamic Characteristics of a Circular Body Earth-to-Orbit Vehicle

NASA Technical Reports Server (NTRS)

Lepsch, Roger A., Jr.; Ware, George M.; MacConochie, Ian O.

1996-01-01

A test of a generic reusable earth-to-orbit transport was conducted in the 7- by 10-Foot high-speed tunnel at the Langley Research Center at Mach number 0.3. The model had a body with a circular cross section and a thick clipped delta wing as the major lifting surface. For directional control, three different vertical fin arrangements were investigated: a conventional aft-mounted center vertical fin, wingtip fins, and a nose-mounted vertical fin. The configuration was longitudinally stable about the estimated center-of-gravity position of 0.72 body length and had sufficient pitch-control authority for stable trim over a wide range of angle of attack, regardless of fin arrangement. The maximum trimmed lift/drag ratio for the aft center-fin configuration was less than 5, whereas the other configurations had values of above 6. The aft center-fin configuration was directionally stable for all angles of attack tested. The wingtip and nose fins were not intended to produce directional stability but to be active controllers for artificial stabilization. Small rolling-moment values resulted from yaw control of the nose fin. Large adverse rolling-moment increments resulted from tip-fin controller deflection above 13 deg angle of attack. Flow visualization indicated that the adverse rolling-moment increments were probably caused by the influence of the deflected tip-fin controller on wing flow separation.

Investigation of configuration effects on entry heating distributions at Mach no. equal 8.0 (OH41). [for wind tunnel model of space shuttle orbiter

NASA Technical Reports Server (NTRS)

Gorowitz, H.; White, R.; Derrico, A.

1973-01-01

Aerodynamic heating data were obtained on 0.006 scale models of four Rockwell International SSV double delta wing Orbiters in the Mach 8 variable density tunnel. A model of two previously tested Rockwell International Orbiters which are identified in the Configuration Description of this report were also tested. Orbiter surfaces were thermally mapped from the laminar through turbulent flight regimes during re-entry. Various modifications were made to model lower surfaces to determine the cause of transition in the vicinity of 3.0 million Reynolds number per foot. Re-entry data were acquired for angles of attack from 25 through 35 degrees at nominal Reynolds numbers per foot of 1.0, 2.0, 2.3, 2.5, 3.0, 3.5, 4.5 and 6.0 million utilizing the phase change paint technique. Launch data were acquired on the model upper surfaces for angles of attack of 0 and -5 degrees at nominal Reynolds numbers per foot of 3.0 and 6.0 million. A total of 70 orbiter heating runs and 6 material sample sphere runs were completed.

Nano-mechanical properties and structural of a 3D-printed biodegradable biomimetic micro air vehicle wing

NASA Astrophysics Data System (ADS)

Salami, E.; Montazer, E.; Ward, T. A.; Ganesan, P. B.

2017-06-01

The biomimetic micro air vehicles (BMAV) are unmanned, micro-scaled aircraft that are bio-inspired from flying organisms to achieve the lift and thrust by flapping their wings. The main objectives of this study are to design a BMAV wing (inspired from the dragonfly) and analyse its nano-mechanical properties. In order to gain insights into the flight mechanics of dragonfly, reverse engineering methods were used to establish three-dimensional geometrical models of the dragonfly wings, so we can make a comparative analysis. Then mechanical test of the real dragonfly wings was performed to provide experimental parameter values for mechanical models in terms of nano-hardness and elastic modulus. The mechanical properties of wings were measured by nanoindentre. Finally, a simplified model was designed and the dragonfly-like wing frame structure was bio-mimicked and fabricated using a 3D printer. Then mechanical test of the BMAV wings was performed to analyse and compare the wings under a variety of simplified load regimes that are concentrated force, uniform line-load and a torque. This work opened up the possibility towards developing an engineering basis for the biomimetic design of BMAV wings.