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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 253 records · Page 14
All sourcesNASA NTRSDOE OSTI
NASA NTRS1959

An Experimental Investigation of a Triangular Wing of Aspect Ratio 2 and a Body Warped to be Trimmed at M = 2.24

A cambered and twisted triangular wing of aspect ratio 2 in combination with a cambered body was investigated experimentally to determine the effectiveness of the camber in reducing the drag due to lift at trim at supersonic speeds. Four arrangements were tested comprising all combinations of a symmetrical and a cambered wing with a symmetrical and a cambered body. The camber shape investigated was derived by linearized lifting surface theory for triangular wings with sonic leading edges and satisfied the requirement that the wing be trimmed at the design Mach number and lift coefficient. The experimental results for the cambered wing and cambered body showed that the drag coefficient at trim was always greater, at the same lift coefficient, than that for the untrimmed symmetrical wing and body. The trim lift coefficient was positive and decreased with increasing Mach number. At the design Mach number of 2.24, the trim lift coefficient was somewhat lower and the drag coefficient was higher than values predicted by linearized lifting surface theory for the wing alone. A comparison of the trim lift-drag ratio of the cambered wing and cambered body with values obtained by trimming the symmetrical wing and symmetrical body either with a canard or a trailing-edge flap showed that, at approximately the design Mach number the cambered configuration developed a somewhat higher value than the trailing-edge flap configuration but a lower value than the canard configuration.

Adams, Gaynor J.↗
NASA NTRS2019

Multi-Objective Optimal Control of the 6-DoF Aeroservoelastic Common Research Model with Aspect Ratio 13.5 Wing

A new 6-DoF aeroservoelastic (ASE) Common Research Model (CRM) provided by The Boeing Company with aspect ratio 13.5 and 17 control surfaces per wing is utilized to demonstrate combined tracking and optimal multi-objective control. The multi-objective controller is derived on the closed loop tracking controller, and utilizes state and gust estimates provided by an extended state observer. Various methods of model reduction useful for control and estimation are presented. A computationally efficient MATLAB/Simulink simulation is presented which includes actuator dynamics, rate and deflection saturation limits, and gust disturbance inputs. The platform is used to demonstrate excellent 6-DoF tracking control performance coupled with the multi-objective controller, which is shown to effectively reduce structural mode movement, wing root bending moment, and drag. State and gust estimation is also shown to perform well, even when derived and/or implemented with significantly fewer states than the original full-sized model.

Drew, Michael C.↗
NASA NTRSDate not supplied

Aeroelastic Simulation of Wind Tunnel Scale Aspect Ratio 13.5 Common Research Model with Trailing Edge Mini-Flaps

This paper presents a computation study of an aspect ratio 13.5 wind tunnel scale Common Research Model (CRM) with trailing edge mini-flaps using FUN3D. The aeroelastic simulations are conducted under free-air and inside wind tunnel conditions. The NASA Langley Research Center Transonic Dynamic Tunnel (TDT) is a closed circuit and slotted-test-section with a 16 ft ×16 ft test section. The structure mode shapes are extracted from the latest updated NASTRAN Finite-Element Model (FEM) model. A dynamic pressure sweep is performed to determine the control reversal boundary of the trailing edge mini-flaps. The wind tunnel wall interference effects on the mini-flap control derivatives are analyzed based on the simulations.

Aeroelastic↗
NASA NTRSDate not supplied

Numerical Simulation of An Aspect Ratio 13.5 Common Research Model with Trailing Edge Mini-Flaps

This paper presents an aeroelastic modeling study of an aspect ratio 13.5 wind tunnel scale Common Research Model (CRM) with trailing edge mini-flaps. FUN3D is used for the aeroelastic steady and unsteady simulations at Mach 0.85 and dynamic pressure 230 psf. The aeroelastic simulation results are compared with the in-house developed VSPAERO aeroelastic model results. The incremental aerodynamic coefficients predicted by FUN3D and VSPAERO are close except when the mini-flap deflection angle is 10˚. The control reversal behavior is observed for the wing tip mini-flap. A series of dynamic pressure sweep simulations are performed to determine the control reversal onset point. Unsteady simulation results show that there is no buffet phenomenon observed at angle of attacks 0˚ and 2˚ when the mini-flap is deflected 10˚. The mini-flap oscillation simulations show the nonlinear behavior appears when the deflection angle is about 4˚.

Mini-flap↗
NASA NTRSDate not supplied

Multi-Objective Optimal Control of High Aspect Ratio Wing Wind Tunnel Model

This paper presents the design of a multi-objective optimal controller for the 13.5 aspect ratio Common Research Model (CRM) wind tunnel model. A Pareto front analysis is performed to provide the optimal tradeoff between the competing objectives - gust load alleviation, maneuver load alleviation, and drag minimization - so that the weighting matrices for the control law are selected optimally. Also presented are two formulations of the controller: a state feedback with Kalman filter design, and an output feedback form where sensors are used directly for feedback control. The controller demonstrates a reduced peak-to-peak strain of 45.78% while also reducing drag and mean bending moment. An output feedback formulation for gust load alleviation is formulated and produces a 53% reduction in the strain response due to a gust input.

Gust Load Alleviation↗
NASA NTRS1987

Navier-Stokes computations of vortical flows over low aspect ratio wings

An upwind-biased finite-volume algorithm is applied to the low-speed flow over a low aspect ratio delta wing from zero to forty degrees angle of attack. The differencing is second-order accurate spatially, and a multigrid algorithm is used to promote convergence to the steady state. The results compare well with the detailed experiments of Hummel (1983) and others for a Re(L) of 0.95 x 10 to the 6th. The predicted maximum lift coefficient of 1.10 at thirty-five degrees angle of attack agrees closely with the measured maximum lift of 1.06 at thirty-three degrees. At forty degrees angle of attack, a bubble type of vortex breakdown is evident in the computations, extending from 0.6 of the root chord to just downstream of the trailing edge.

Thomas, J. L.↗
NASA NTRS1948

Aerodynamic Characteristics at Subsonic and Supersonic Mach Numbers of a Thin Triangular Wing of Aspect Ratio 2 II : Maximum Thickness at Midchord

The lift, drag, and pitching-moment characteristics of a triangular wing, having an aspect ratio of 2 and a symmetrical double-wedge profile of 5-percent-chord maximum thickness at midchord, have been evaluated from wind-tunnel tests at Mach numbers from 0.50 to 0.975 and from 1.09 to 1.49 and at Reynolds numbers ranging from 0.67 to 0.85 million. The lift, drag, and pitching-moment coefficients of the triangular wing with a leading-edge sweepback of approximately 63 degrees did not exhibit the irregular variations with Mach number at high subsonic and low supersonic Mach numbers that are characteristic of unswept wings. The lift-curve slope increased steadily with Mach number below unity and declined slowly beyond the Mach number of 1.13. A substantial rise in the minimum drag coefficient occurred between Mach numbers of 0.95 and 1.20 with an associated reduction in the maximum lift-drag ratio. The aerodynamic center shifted rearward toward the centroid of area of the wing with increasing Mach number below 0.975; whereas above 1.09 it coincided with the centroid.

Walker, Harold J↗
NASA NTRS1953

Aerodynamic characteristics of low-aspect-ratio wings at high supersonic Mach numbers

This paper presents some recently obtained data on the aerodynamic characteristics of low-aspect-ratio wings at supersonic Mach numbers of 4.04 and 6.9 and discusses some new methods of predicting the lift and drag of such wings. Data on lifting wings in the Mach number range above 2.5 are not plentiful and most of the available data may be found in NACA RM L51D17, NACA RM L51D30, NACA RM L52D15a, NACA RM L52K19, NACA RM L53D03, and NACA RM L53D30a

Ulmann, Edward F↗
NASA NTRS1940

Investigation of the Characteristics of a High-Aspect-Ratio Wing in the Langley 8-Foot High-Speed Tunnel

An investigation of the characteristics of a wing with an aspect ratio of 9.0 and an NACA 65-210 airfoil section has been made at Mach number up to 0.925. The wing tested has a taper ratio of 2.5:1.0, no twist, dihedral, or sweepback, and 20-percent - chord 37.5-percent-semispan plain ailerons. The results showed that serious changes in the normal-force characteristics occurred when the Mach number was increased above 0.74 at angles of attack between 4 deg. and 10 deg. and above 0.80 at 0 deg. angle of attack.Because of small outboard shifts in the lateral center of load, the bending moment at the root for conditions corresponding to a 3g pull-out at an altitude of 35,000 feet increased by approximately 5% when the Much number was increased beyond 0.83 the negative pitching moments for the high angles of attack increased, whereas those for the low angles of attack decreased with a resulting large increase in the negative slope of the pitching-moment curves. A large increase occurred in the values of the drag coefficients for the range of lift coefficients needed for level flight at an altitude of 35,000 feet when the Mach number was increased beyond a value of 0.80. The wakes at a station 2.82 root chords behind the wing quarter-chord line extended approximately a chord above the wing chord line for the angles of attack required to recover from high-speed dives at high Mach numbers.

Whitcomb, Richard T.↗
DOE OSTI2026

Initial electron cyclotron heating/current drive scoping study for the SMall Aspect Ratio Tokamak (SMART)

In this work an initial electron cyclotron heating/current drive (EC H&CD) scoping study is presented for the SMall Aspect Ratio Tokamak (SMART), constructed and operated by the Plasma Science and Fusion Technology (PSFT) Laboratory of the University of Seville. We consider two planned phases of operation with different magnitude of magnetic field (B 0.4, and 1 T). This EC H&CD scoping study is carried out by using the ray tracing code TRAVIS, which allows to quickly perform a series of scans on the main SMART plasma scenarios. A scan in the poloidal and toroidal launching angles is performed for the B 0.4 T plasma scenario assuming three launching locations (mid-plane, off-mid-plane, and top-launcher) and the extraordinary (X-) mode polarization. The absorption occurs at the second harmonic and, overall, the highest power absorption is obtained launching the EC beam from the top of the machine. A numerical analysis of the fundamental X-mode EC H&CD start-up regime for the B 1.0 T plasma scenario with two different wave frequencies (28 and 35 GHz) is presented showing a very high current drive efficiency.

Bertelli, Nicola [Princeton Plasma Physics Laborat↗
NASA NTRS1985

Lift due to thickness for low aspect ratio wings in incompressible flow

The problem under consideration is a numerical study of the effects of thickness on lift for low aspect ratio wings in steady incompressible inviscid flow at moderate angles of attack. At these angles of attack the flow separates along the leading edge giving rise to a lift substantially higher than that computed by classical attached flow potential theory. The problem is treated as a perturbation expansion in a small thickness parameter. The lifting elements of the flow are modeled using a nonlinear vortex lattice method which replaces the leading and trailing edge vortex sheets by segmented straight vortex filaments. The thickness elements of the flow are modeled with a mean plane source distribution and a modification to the wing boundary conditions. Results are obtained for wings with biconvex and NACA 0012 sections which compare well with available experimental data. The important observation that the effect of thickness is to decrease the lift is made.

Dodbele, S. S.↗
NASA NTRS1933

Aerodynamic tests of a low aspect ratio tapered wing with various flaps, for use on tailless airplanes

Wind tunnel tests were made of a model wing having an aspect ratio of 3 and a tapered plan form with a straight trailing edge. The model had the Clark Y airfoil section throughout it's entire span and had no washout, depending on a trailing-edge flap for longitudinal balance and control. The flap had a constant chord and was divided into four equal portions along the span. The tests were made with the entire flap deflected to obtain longitudinal control and balance, and also with the inner portions deflected alone, and with the outer portions deflected alone. It was found that the simple wing with no washout or change of basic section along the span has aerodynamic characteristics well suited for use on tailless airplanes. A higher lift coefficient was obtained with the full-span flap deflected as a unit to give longitudinal balance than was obtained with either the inner or the outer portions of the flap deflected.

Weick, Fred E↗
NASA NTRS1996

Validation of High Aspect Ratio Cooling in a 89 kN (20,000 lb(sub f)) Thrust Combustion Chamber

In order to validate the benefits of high aspect ratio cooling channels in a large scale rocket combustion chamber, a high pressure, 89 kN (20,000 lbf) thrust, contoured combustion chamber was tested in the NASA Lewis Research Center Rocket Engine Test Facility. The combustion chamber was tested at chamber pressures from 5.5 to 11.0 MPa (800-1600 psia). The propellants were gaseous hydrogen and liquid oxygen at a nominal mixture ratio of six, and liquid hydrogen was used as the coolant. The combustion chamber was extensively instrumented with 30 backside skin thermocouples, 9 coolant channel rib thermocouples, and 10 coolant channel pressure taps. A total of 29 thermal cycles, each with one second of steady state combustion, were completed on the chamber. For 25 thermal cycles, the coolant mass flow rate was equal to the fuel mass flow rate. During the remaining four thermal cycles, the coolant mass flow rate was progressively reduced by 5, 6, 11, and 20 percent. Computer analysis agreed with coolant channel rib thermocouples within an average of 9 percent and with coolant channel pressure drops within an average of 20 percent. Hot-gas-side wall temperatures of the chamber showed up to 25 percent reduction, in the throat region, over that of a conventionally cooled combustion chamber. Reducing coolant mass flow yielded a reduction of up to 27 percent of the coolant pressure drop from that of a full flow case, while still maintaining up to a 13 percent reduction in a hot-gas-side wall temperature from that of a conventionally cooled combustion chamber.

Wadel, Mary F.↗
NASA NTRS1978

A theoretical investigation of the aerodynamics of low-aspect-ratio wings with partial leading-edge separation

A numerical method is developed to predict distributed and total aerodynamic characteristics for low aspect-ratio wings with partial leading-edge separation. 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 mid-points 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. Comparison of the predicted results with complete leading-edge separation has shown reasonably good agreement. For cases with partial leading-edge separation, the lift is found to be highly nonlinear with angle of attack.

Mehrotra, S. C.↗
NASA NTRS1985

Detailed pressure distribution measurements obtained on several configurations of an aspect-ratio-7 variable twist wing

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.

Holbrook, G. T.↗
NASA NTRS1987

Loss of lift due to thickness for low-aspect-ratio wings in incompressible flow

The problem under consideration is a numerical study of the effects of thickness on lift for low-aspect-ratio wings in steady incompressible inviscid flow at moderate angles of attack. At these angles of attack the flow separates along the leading edge giving rise to a lift substantially higher than that computed by classical attached-flow potential theory. The problem is treated as a perturbation expansion in a small thickness parameter. The lifting elements of the flow are modeled using a nonlinear vortex-lattice method which replaces the leading and trailing-edge vortex sheets by segmented straight vortex filaments. The thickness elements of the flow are modeled with a mean-plane source distribution and a modification to the wing boundary conditions. Results are obtained for wings with biconvex and NACA 0012 sections which compare well with available experimental data. The important observation that the effect of thickness is to decrease the lift is made.

Dodbele, S. S.↗