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At least 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Skylon Aerodynamics and SABRE Plumes

An independent partial assessment is provided of the technical viability of the Skylon aerospace plane concept, developed by Reaction Engines Limited (REL). The objectives are to verify REL's engineering estimates of airframe aerodynamics during powered flight and to assess the impact of Synergetic Air-Breathing Rocket Engine (SABRE) plumes on the aft fuselage. Pressure lift and drag coefficients derived from simulations conducted with Euler equations for unpowered flight compare very well with those REL computed with engineering methods. The REL coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5, because the engineering estimates did not account for the increasing favorable (in terms of drag and lift coefficients) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown−a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plumeinduced flow separation are other potential risks. The development of an operational reusable launcher from the Skylon concept necessitates the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments.

Mehta, Unmeel↗
NASA NTRS1943

The Effect of Compressibility on the Growth of the Laminar Boundary Layer on Low-Drag Wings and Bodies

The development of the laminar boundary layer in a compressible fluid is considered. Formulas are given for determining the boundary-layer thickness and the ratio of the boundary-layer Reynolds number to the body Reynolds number for airfoils and bodies of revolution. It i s shown that the effect of compressibility will profoundly alter the Reynolds number corresponding to the upper limit of the range of the low-drag coefficients . The available data indicate that for low-drag and high critical compressibility speed airfoils and bodies of revolution, this effect is favorable.

Allen, H. Julian↗
NASA NTRS2018

Steady and Unsteady Excitation of Separated Flow over the NASA Hump Model

Separated flow and its control over the NASA hump model were investigated at subsonic speeds. Three-dimensional, unsteady fluid dynamic simulations were supplemented by wind tunnel measurements. Flow control was implemented by means of spatially distributed discrete jets operating in steady and unsteady modes. The effects of excitation amplitude and frequency were studied numerically and experimentally. Several integral parameters were explored as quality metrics. In addition to the existing ones, two new integral parameters, which are slightly modified versions of the normal force and moment coefficients, were introduced. These two parameters together with pressure drag coefficient were found to be well correlated with the flow control and used in the performance evaluation of different flow control methods including, zero net mass flux actuators, steady suction, sweeping jet actuators, and the currently studied steady and unsteady excitations. For the cases tested, it was found that the unsteady excitation is superior to the steady excitation and slightly better than the sweeping jet actuators whereas the steady suction was found to be the most effective. Although the numerical simulations overpredict the separation bubble, these simulations capture the salient features of flow separation control and hence help us to understand the effect of steady/unsteady excitation on the separated flow.

Koklu, Mehti↗
NASA NTRSDate not supplied

Aerodynamic Design Optimization for Natural Laminar Flow Airfoils

Natural laminar flow technology is a passive laminar flow control (LFC) strategy that seeks to delay the onset of boundary-layer transition (BLT) through shape optimization to reduce the drag of the aerodynamic vehicle. Adjoint-based design optimization for LFC is proposed in an integrated multidisciplinary framework, which includes the computational fluid dynamics (CFD), geometry and grid deformation, and linear stability analysis (LSA) for transition prediction. In particular, the BLT location is predicted using the dual N-factor method that is based on a linear stability theory (LST) eigenvalue problem. The dual N-factor criterion accounts for the amplification of planar Tollmien-Schlichting (TS) and stationary crossflow (CF) boundary-layer instabilities to predict the transition location in three-dimensional boundary-layer flows. The adjoint-based shape optimization procedure is based on an iteratively coupled CFD and LSA methodology to converge the transition location and flow solutions, as well as to calculate the sensitivities of the aerodynamic metrics of interest with respect to the flow and shape design parameters. The RAE 2822 airfoil at 0 and 30 degrees yaw angles, an angle of attack of 0.72 degrees, and subsonic conditions (M∞ = 0.19, Rec = 5.6 × 106 ) are used as baseline configurations for design optimization. The angle of attack and the vertical displacement of free-form-deformation control points are used as design variables to reduce the drag coefficient while reaching a specified lift coefficient. The optimized unswept airfoil designs achieve a 30% drag reduction accompanied by a downstream shift of the transition locations over both suction and pressure sides of the airfoil. The initial design iterations for the swept case also show a favorable trend in the drag reduction with transition delay over both sides.

Transition↗
NASA NTRS2015

Semi-Empirical Prediction of Aircraft Low-Speed Aerodynamic Characteristics

This paper lays out a comprehensive methodology for computing a low-speed, high-lift polar, without requiring additional details about the aircraft design beyond what is typically available at the conceptual design stage. Introducing low-order, physics-based aerodynamic analyses allows the methodology to be more applicable to unconventional aircraft concepts than traditional, fully-empirical methods. The methodology uses empirical relationships for flap lift effectiveness, chord extension, drag-coefficient increment and maximum lift coefficient of various types of flap systems as a function of flap deflection, and combines these increments with the characteristics of the unflapped airfoils. Once the aerodynamic characteristics of the flapped sections are known, a vortex-lattice analysis calculates the three-dimensional lift, drag and moment coefficients of the whole aircraft configuration. This paper details the results of two validation cases: a supercritical airfoil model with several types of flaps; and a 12-foot, full-span aircraft model with slats and double-slotted flaps.

Olson, Erik D.↗
NASA NTRS1977

Application of numerical optimization to the design of supercritical airfoils without drag-creep

Recent applications of numerical optimization to the design of advanced airfoils for transonic aircraft have shown that low-drag sections can be developed for a given design Mach number without an accompanying drag increase at lower Mach numbers. This is achieved by imposing a constraint on the drag coefficient at an off-design Mach number while the drag at the design Mach number is the objective function. Such a procedure doubles the computation time over that for single design-point problems, but the final result is worth the increased cost of computation. The ability to treat such multiple design-point problems by numerical optimization has been enhanced by the development of improved airfoil shape functions. Such functions permit a considerable increase in the range of profiles attainable during the optimization process.

Hicks, R. M.↗
NASA NTRS1992

Vortex trapping on a 60 degree delta wing

An experiment was conducted to examine the vortex trapping concept on a 60 deg sweep delta wing at low angles of attack. Using tapered leading edge fences, stable vortices have been successfully trapped. The resulting increase in lift coefficient can be as high as 0.4, which amounts to a 6.5 deg shift in the lift coefficient curve of the delta wing. The corresponding increase in drag coefficient, however, is also significant.

Buchholz, Mark D.↗
NASA NTRS1994

Off-Axis Drag of Dendrite Fragments at Low Reynolds Number

The aim of the present investigation is to characterize the motion of dendrite fragments falling under the influence of gravity in a uniform liquid medium at low Reynolds number. In an earlier study, Zakhem, Weidman and de Groh (1992) reported on the settling speed of model equiaxed dendrite grains released along their axis of symmetry. In this follow-up study uniaxial model dendrite grains were released off-axis to observe and document their motion at different orientations. It was hypothesized that the dendrite models might rotate when released off-axis in which case an attempt would be made to document the ensuing unsteady motion. This latter event turned out to be in fact true: at the small but finite Reynolds numbers that existed, each uniaxial dendrite slowly rotated towards its equilibrium orientation while failing under the influence of gravity. In addition to completing the original goal, we have made use of a beads-on-a shell Stokes flow code to numerically determine the drag coefficient for capsules, i.e.. uniaxial dendrites without arms. The drag on horizontally and vertically falling capsules are reported and compared with measurements.

Weidman, P. D.↗
NASA NTRSDate not supplied

USM3D-ME Contributions to the 5th AIAA High Lift Prediction Workshop

This paper presents the results of Reynolds-averaged Navier-Stokes (RANS) simulations conducted by NASA’s flow solver, mixed-element USM3D (USM3D-ME), for the 5th AIAA High-Lift Prediction Workshop. As part of the Fixed-Grid RANS Technology Focus Group (TFG), these simulations were performed to assess the accuracy and efficiency of the USM3D-ME solutions in predicting high-lift flows. The High-Lift Common Research Model (CRM-HL) served as the primary geometry. Several CRM-HL configurations were used for three case studies: a verification study (Case 1), a configuration buildup study (Case 2), and a Reynolds-number variation study (Case 3). Overall, USM3D-ME RANS results aligned with the solutions selected by the Fixed-Grid RANS TFG and available wind tunnel data Simulations for Cases 1 and Configuration 2.1 achieved machine-zero residual convergence, with aerodynamic coefficients converging to steady-state values. However, Configurations 2.2-2.4 and Case 3 encountered iterative- and grid-convergence challenges, particularly at high angles of attack. Compared with the experimental data available for Configurations 2.2-2.4, close agreement was demonstrated at low angles of attack. However, for angles of attack approaching the maximum lift conditions, the predicted lift coefficient and pitching moment deviated from experimental values. The drag-coefficient predictions were in a relatively good agreement, however, slight overpredictions were observed at the highest angle of attack corresponding to the maximum-lift condition. Although iterative convergence for Configurations 2.2-2.4 at high angles of attack remains a persistent challenge, averaging aerodynamic coefficients over the last 5000 iterations yielded satisfactory agreement with the available wind tunnel experimental data. During the workshop, the lack of iterative convergence was attributed to the vortex structures emanating from the slat brackets. To investigate this issue further, post-workshop simulations were conducted on Configuration 2.2. In one study, RANS simulations were performed on a simplified geometry with the slat brackets removed. The second study focused on performing URANS simulations on the original Configuration 2.2 geometry. Preliminary results from both studies are presented and compared with wind tunnel data for Configuration 2.2. Consistent with the findings of other participants in the Fixed-Grid RANS TFG, this study emphasizes the necessity for further exploration and advancement in RANS technology for predicting high-lift flows.

CFD↗
NASA NTRSDate not supplied

USM3D-ME Contributions to the 5th AIAA High Lift Prediction Workshop

This paper presents the results of Reynolds-averaged Navier-Stokes (RANS) simulations conducted by the NASA flow solver, mixed-element USM3D (USM3D-ME), for the 5th AIAA High-Lift Prediction Workshop. As part of the Fixed-Grid RANS Technology Focus Group (TFG), these simulations were performed to assess the accuracy and efficiency of the USM3D-ME solutions in predicting high-lift flows. The High-Lift Common Research Model (CRM-HL) served as the primary geometry. Several CRM-HL configurations were used for three case studies: a verification study (Case 1), a configuration buildup study (Case 2), and a Reynolds-number variation study (Case 3). Overall, USM3D-ME RANS results aligned with the solutions selected by the Fixed-Grid RANS TFG and available wind tunnel data. Simulations for Case 1 and Configuration 2.1 achieved machine-zero residual convergence, with aerodynamic coefficients converging to steady-state values. However, Configurations 2.2-2.4 and Case 3 encountered iterative- and grid-convergence challenges, particularly at high angles of attack. Compared with the experimental data available for Configurations 2.2-2.4, close agreement was demonstrated at low angles of attack. However, for angles of attack approaching the maximum lift conditions, the predicted lift coefficient and pitching moment deviated from experimental values. The drag-coefficient predictions were in relatively good agreement, however, slight overpredictions were observed at the highest angle of attack corresponding to the maximum-lift condition. Although iterative convergence for Configurations 2.2-2.4 at high angles of attack remains a persistent challenge, averaging aerodynamic coefficients over the last 5000 iterations yielded satisfactory agreement with the available wind tunnel experimental data. During the workshop, the lack of iterative convergence was attributed to the vortex structures emanating from the slat brackets. To investigate this issue further, post-workshop simulations were conducted on Configuration 2.2. In one study, RANS simulations were performed on a simplified geometry with the slat brackets removed. The second study focused on performing unsteady RANS (URANS) simulations on the original Configuration 2.2 geometry. Preliminary results from both studies are presented and compared with wind tunnel data for Configuration 2.2. Consistent with the findings of other participants in the Fixed-Grid RANS TFG, this study emphasizes the necessity for further exploration and advancement in RANS technology for predicting high-lift flows.

Aerodynamics↗
DOE OSTI2024

Misaligned Wind‐Waves Behind Atmospheric Cold Fronts

Abstract Atmospheric fronts embedded in extratropical cyclones are high‐impact weather phenomena, contributing significantly to mid‐latitude winter precipitation. The three vital characteristics of the atmospheric fronts, high wind speeds, abrupt change in wind direction, and rapid translation, force the induced surface waves to be misaligned with winds exclusively behind the cold fronts. The effects of the misaligned waves under atmospheric cold fronts on air‐sea fluxes remain undocumented. Using the multi‐year in situ near‐surface observations and direct covariance flux measurements from the Pioneer Array off the coast of New England, we find that the majority of the passing cold fronts generate misaligned waves behind the cold front. Once generated, the waves remain misaligned, on average, for about 8 hr. The parameterized effect of misaligned waves in a fully coupled model significantly increases the roughness length (185%), drag coefficient (19%), and air‐sea momentum flux (11%). The increased surface drag reduces the wind speeds in the surface layer. The upward turbulent heat flux is weakly decreased by the misaligned waves because of the decrease in temperature and humidity scaling parameters being greater than the increase in friction velocity. The misaligned wave effect is not accurately represented in a commonly used wave‐based bulk flux algorithm. Yet, considering this effect in the current formulation improves the overall accuracy of parameterized momentum flux estimates. The results imply that better representing a directional wind‐wave coupling in the bulk formula of the numerical models may help improve the air‐sea interaction simulations under the passing atmospheric fronts in the mid‐latitudes.

Oceanography↗
NASA NTRS1960

Hydrodynamic and Aerodynamic Characteristics of a Model of a Supersonic Multijet Water-Based Aircraft Equipped with Supercavitating Hydrofoils

The hydrodynamic and aerodynamic characteristics of a model of a multijet water-based Mach 2.0 aircraft equipped with hydrofoils have been determined. Takeoff stability and spray characteristics were very good, and sufficient excess thrust was available for takeoff in approximately 32 seconds and 4,700 feet at a gross weight of 225,000 pounds. Longitudinal and lateral stability during smooth-water landings were good. Lateral stability was good during rough-water landings, but forward location of the hydrofoils or added pitch damping was required to prevent diving. Hydrofoils were found to increase the aerodynamic lift-curve slope and to increase the aerodynamic drag coefficient in the transonic speed range, and the maximum lift-drag ratio decreased from 7.6 to 7.2 at the cruise Mach number of 0.9. The hydrofoils provided an increment of positive pitching moment over the Mach number range of the tests (0.6 to 1.42) and reduced the effective dihedral and directional stability.

McKann, Robert E.↗
NASA NTRS1955

Longitudinal Aerodynamic Characteristics and Effect of Rocket Jet on Drag of Models of the Hermes A-3A and A-3B Missiles in Free Flight at Mach Numbers From 0.6 to 2.0

A free-flight investigation over a Mach number range from 0.6 to 2.0 has been conducted to determine the longitudinal aerodynamic characteristics and effect of rocket jet on zero-lift drag of 1/5-scale models of two ballistic-type missiles, the Hermes A-3A and A-3B. Models of both types of missiles exhibited very nearly linear normal forces and pitching moments over the angle-of-attack range of 8 deg to -4 deg and Mach number range tested. The centers of pressure for both missiles were not appreciably affected by Mach number over the subsonic range; however, between a Mach number of 1.02 and 1.50 the center of pressure for the A-3A model moved forward 0.34 caliber with increasing Mach number. At a trim angle-of-attack of approximately 30 deg, the A-3A model indicated a total drag coefficient 30% higher than the power-off zero-lift drag over the subsonic Mach number range and 10% higher over the supersonic range. Under the conditions of the present test, and excluding the effect of the jet on base drag, there was no indicated effect of the propulsive jet on the total drag of the A-3A model. The propulsive jet operating at a jet pressure ratio p(sub j)/p(sub o) of 0.8 caused approximately 100% increase in base drag over the Mach number range M = 0.6 to 1.0. This increase in base drag amounts to 15% of the total drag. An underexpanded jet operating at jet pressure ratios corresponding approximately to those of the full-scale missile caused a 22% reduction in base drag at M = 1.55 (p(sub j)/p(sub o) = 1.76) but indicated no change at M = 1.30 (p(sub j)/p(sub o) = 1.43). At M = 1.1 and p(sub j)/p(sub o) = 1.55, the jet caused a 50% increase in base drag.

Jackson, H. Herbert↗
NASA NTRS1961

Effect of Groundboard Height on the Aerodynamic Characteristics of a Lifting Circular Cylinder Using Tangential Blowing from Surface Slots for Lift Generation

A wind-tunnel investigation has been made to determine the ground effect on the aerodynamic characteristics of a lifting circular cylinder using tangential blowing from surface slots to generate high lift coefficients. The tests were made on a semispan model having a length 4 times the cylinder diameter and an end plate of 2.5 diameters. The tests were made at low speeds at a Reynolds number of approximately 290,000, over a range of momentum coefficients from 0.14 to 4.60, and over a range of groundboard heights from 1.5 to 10 cylinder diameters. The investigation showed an earlier stall angle and a large loss of lift coefficient as the groundboard was brought close to the cylinder when large lift coefficients were being generated. For example, at a momentum coefficient of 4.60 the maximum lift coefficient was reduced from a value of 20.3 at a groundboard height of 10 cylinder diameters to a value of 8.7 at a groundboard height of 1.5 cylinder diameters. In contrast to this there was little effect on the lift characteristics of changes in groundboard height when lift coefficients of about 4.5 were being generated. At a height of 1.5 cylinder diameters the drag coefficients generally increased rapidly when the slot position angle for maximum lift was exceeded. Slightly below the slot position angle for maximum lift, the groundboard had a beneficial effect, that is, the drag for a given lift was less near the groundboard than away from the groundboard. The variation of maximum circulation lift coefficient (maximum lift coefficient minus momentum coefficient) obtained in this investigation is in general agreement with a theory developed for a jet-flap wing which assumes that the loss in circulation is the result of blockage of the main stream beneath the wing.

Lockwood, Vernard E.↗
DOE OSTI2026

OC7 phase I: CFD investigation of viscous forces on rectangular members of semisubmersibles

In Phase I of the Offshore Code Comparison Collaboration 7 (OC7) project, we aim to address key questions related to the modeling of viscous hydrodynamic loads on floating platforms. Here, this paper presents results from a collaborative effort within Phase I of the OC7 project. We investigate viscous forces acting on the rectangular members of semisubmersible platforms using high-fidelity computational fluid dynamics (CFD) simulations. The parametric CFD analyses include two-and-a-half-dimensional (2.5D) simulations of pontoon sections of two different semisubmersible platform designs (the VolturnUS-S and the INO WINDMOOR) and full three-dimensional (3D) structure simulations of the VolturnUS-S platform. We extract both the added mass and drag coefficients that are typically used in engineering-level hydrodynamic models from the 2.5D forced oscillation simulations with various Keulegan-Carpenter numbers and compare them to the coefficients from the 3D forced oscillation simulations. We observe consistent variation of the coefficients along the pontoon legs due to 3D platform and free-surface effects. With the breadth of the simulations and component-level detailed analyses, this work establishes a useful database and several empirical formulas for selecting hydrodynamic coefficients when modeling semisubmersible platforms in engineering tools.

13 HYDRO ENERGY↗
NASA NTRS2004

Calculated Drag of an Aerial Refueling Assembly Through Airplane Performance Analysis

The aerodynamic drag of an aerial refueling assembly was calculated during the Automated Aerial Refueling project at the NASA Dryden Flight Research Center. An F/A-18A airplane was specially instrumented to obtain accurate fuel flow measurements and to determine engine thrust. A standard Navy air refueling store with a retractable refueling hose and paradrogue was mounted to the centerline pylon of the F/A-18A airplane. As the paradrogue assembly was deployed and stowed, changes in the calculated thrust of the airplane occurred and were equated to changes in vehicle drag. These drag changes were attributable to the drag of the paradrogue assembly. The drag of the paradrogue assembly was determined to range from 200 to 450 lbf at airspeeds from 170 to 250 KIAS. Analysis of the drag data resulted in a single drag coefficient of 0.0056 for the paradrogue assembly that adequately matched the calculated drag for all flight conditions. The drag relief provided to the tanker airplane when a receiver airplane engaged the paradrogue is also documented from 35 to 270 lbf at the various flight conditions tested. The results support the development of accurate aerodynamic models to be used in refueling simulations and control laws for fully autonomous refueling.

Vachon, Michael Jacob↗
NASA NTRS1947

Development of Inboard Nacelle for the XB-36 Airplane

A series of investigations of several 1/14-scale models of an inboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels. The purpose of these investigations was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. As a result of these investigations, a nacelle was developed which had external drag coefficients considerably lower than the original basic form with the external nacelle drag approximately one-half to two-thirds of those of conventional tractor designs. The largest reductions in drag resulted from sealing the gaps between the wing flaps and nacelle, reducing the thickness of the nacelle training-edge lip, and bringing the under-wing air inlet to the wing leading edge. It was found that without the engine cooling fan adequate cooling air would be available for all conditions of flight except for cruise and climb at 40,000 feet. Sufficient oil cooling at an altitude of 40,000 feet may be obtained by the use of flap-type exit doors.

Nuber, Robert J.↗
NASA NTRS1953

Flight Test Results of Rocket-Propelled Buffet-Research Models Having 45 Degree Sweptback Wings and 45 Degree Sweptback Tails Located in the Wing Chord Plane

Three rocket-propelled buffet-research models have been flight tested to determine the buffeting characteristics of a swept-wing- airplane configuration with the horizontal tail operating near the wing wake. The models consisted of parabolic bodies having 45deg sweptback wings of aspect ratio 3.56, at aspect ratio of 0.3, NACA 64A007 airfoil sections, and tail surfaces of geometry and section identical to the wings. Two tests were conducted with the horizontal tail located in the wing chord plane with fixed incidence angles of -1.5deg on one model and 0deg on the other model. The third test was conducted with no horizontal tail. Results of these tests are presented as incremental accelerations in the body due to buffeting, trim angles of attack, trim normal- and side-force coefficients, wing-tip helix angles, static-directional-stability derivatives , and drag coefficients plotted against Mach number. These data indicate that mild low-lift buffeting was experienced by all models over a range of Mach number from approximately 0.7 to 1.4. It is further indicated that this buffeting was probably induced by wing-body interference and was amplified at transonic speeds by the horizontal tail operating in the wing wake. A longitudinal trim change was encountered by the tail-on models at transonic speeds, but no large changes in side force and no wing dropping were indicated.

Mason, Homer P.↗