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At least 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Development of Multiobjective Optimization Techniques for Sonic Boom Minimization

A discrete, semi-analytical sensitivity analysis procedure has been developed for calculating aerodynamic design sensitivities. The sensitivities of the flow variables and the grid coordinates are numerically calculated using direct differentiation of the respective discretized governing equations. The sensitivity analysis techniques are adapted within a parabolized Navier Stokes equations solver. Aerodynamic design sensitivities for high speed wing-body configurations are calculated using the semi-analytical sensitivity analysis procedures. Representative results obtained compare well with those obtained using the finite difference approach and establish the computational efficiency and accuracy of the semi-analytical procedures. Multidisciplinary design optimization procedures have been developed for aerospace applications namely, gas turbine blades and high speed wing-body configurations. In complex applications, the coupled optimization problems are decomposed into sublevels using multilevel decomposition techniques. In cases with multiple objective functions, formal multiobjective formulation such as the Kreisselmeier-Steinhauser function approach and the modified global criteria approach have been used. Nonlinear programming techniques for continuous design variables and a hybrid optimization technique, based on a simulated annealing algorithm, for discrete design variables have been used for solving the optimization problems. The optimization procedure for gas turbine blades improves the aerodynamic and heat transfer characteristics of the blades. The two-dimensional, blade-to-blade aerodynamic analysis is performed using a panel code. The blade heat transfer analysis is performed using an in-house developed finite element procedure. The optimization procedure yields blade shapes with significantly improved velocity and temperature distributions. The multidisciplinary design optimization procedures for high speed wing-body configurations simultaneously improve the aerodynamic, the sonic boom and the structural characteristics of the aircraft. The flow solution is obtained using a comprehensive parabolized Navier Stokes solver. Sonic boom analysis is performed using an extrapolation procedure. The aircraft wing load carrying member is modeled as either an isotropic or a composite box beam. The isotropic box beam is analyzed using thin wall theory. The composite box beam is analyzed using a finite element procedure. The developed optimization procedures yield significant improvements in all the performance criteria and provide interesting design trade-offs. The semi-analytical sensitivity analysis techniques offer significant computational savings and allow the use of comprehensive analysis procedures within design optimization studies.

Chattopadhyay, Aditi↗
NASA NTRS1959

Investigation at Mach Numbers of 0.20 to 3.50 of a Blended Diamond Wing and Body Combination of Sonic Design but with Low Wave-Drag Increase with Increasing Mach Number

A diamond wing and body combination was designed to have an area distribution which would result in near optimum zero-lift wave-drag coefficients at a Mach number of 1.00, and decreasing wave-drag coefficient with increasing Mach number up to near sonic leading-edge conditions for the wing. The airfoil section were computed by varying their shape along with the body radii (blending process) to match the selected area distribution and the given plan form. The exposed wing section had an average maximum thickness of about 3 percent of the local chords, and the maximum thickness of the center-line chord was 5.49 percent. The wing had an aspect ratio of 2 and a leading-edge sweep of 45 deg. Test data were obtained throughout the Mach number range from 0.20 to 3.50 at Reynolds numbers based on the mean aerodynamic chord of roughly 6,000,000 to 9,000,000. The zero-lift wave-drag coefficients of the diamond model satisfied the design objectives and were equal to the low values for the Mach number 1.00 equivalent body up to the limit of the transonic tests. From the peak drag coefficient near M = 1.00 there was a gradual decrease in wave-drag coefficient up to M = 1.20. Above sonic leading-edge conditions of the wing there was a rise in the wave-drag coefficient which was attributed in part to the body contouring as well as to the wing geometry. The diamond model had good lift characteristics, in spite of the prediction from low-aspect-ratio theory that the rear half of the diamond wing would carry little lift. The experimental lift-curve slope obtained at supersonic speeds were equal to or greater than the values predicted by linear theory. Similarly the other basic aerodynamic parameters, aerodynamic center position, and maximum lift-drag ratios were satisfactorily predicted at supersonic speeds.

Holdaway, George H.↗
NASA NTRS1965

Prediction of Airplane Sonic-Boom Pressure Fields

This paper presents a discussion of the sensitivity of supersonic-transport design and operation to sonic-boom considerations and shows the necessity for a study of these problems early in the development program. Methods of predicting pressure signatures are outlined and examples of the correlation of these estimates with wind-tunnel and flight measurements are shown. Estimates of sonic-boom characteristics for a representative supersonic transport show that in the critical transonic acceleration portion of the flight, overpressures somewhat lower than estimated by the use of far-field assumptions may be expected. Promising design possibilities for the achievement of further overpressure reductions are explored.

Carlson, Harry W.↗
NASA NTRS1995

Effects of Sonic Line Transition on Aerothermodynamics of the Mars Pathfinder Probe

Flow field solutions over the Mars Pathfinder Probe spanning the trajectory through the Martian atmosphere at angles of attack from 0 to 11 degrees are obtained. Aerodynamic coefficients derived from these solutions reveal two regions where the derivative of pitching moment with respect to angle of attack is positive at small angles of attack. The behavior is associated with the transition of the sonic line location between the blunted nose and the windside shoulder of the 70 degree half-angle cone in a gas with a low effective ratio of specific heats. The transition first occurs as the shock layer gas chemistry evolves from highly nonequilibrium to near equilibrium, above approximately 6.5 km/s and 40 km altitude, causing the effective specific heat ratio to decrease. The transition next occurs in an equilibrium flow regime as velocities decrease through 3.5 km/s and the specific heat ratio increases again with decreasing enthalpy. The effects of the expansion over the shoulder into the wake are more strongly felt on the fustrum when the sonic line sits on the shoulder. The transition also produces a counter-intuitive trend in which windside heating levels decrease with increasing angle of attack resulting from an increase in the effective radius of curvature. Six-degree-of-freedom trajectory analyses utilizing the computed aerodynamic coefficients predict a moderate, 3 to 4 degree increase in total angle of attack as the probe, spinning at approximately 2 revolutions per minute, passes through these regions.

Gnoffo, Peter A.↗
NASA NTRS2005

Method for Estimating the Sonic-Boom Characteristics of Lifting Canard-Wing Aircraft Concepts

A method for estimating the sonic-boom overpressures from a conceptual aircraft where the lift is carried by both a canard and a wing during supersonic cruise is presented and discussed. Computer codes used for the prediction of the aerodynamic performance of the wing, the canard-wing interference, the nacelle-wing interference, and the sonic-boom overpressures are identified and discussed as the procedures in the method are discussed. A canard-wing supersonic-cruise concept was used as an example to demonstrate the application of the method.

Mack, Robert J.↗
NASA NTRS2008

Hybrid LES/RANS Simulation of Transverse Sonic Injection into a Mach 2 Flow

A computational study of transverse sonic injection of air and helium into a Mach 1.98 cross-flow is presented. A hybrid large-eddy simulation / Reynolds-averaged Navier-Stokes (LES/RANS) turbulence model is used, with the two-equation Menter baseline (Menter-BSL) closure for the RANS part of the flow and a Smagorinsky-type model for the LES part of the flow. A time-dependent blending function, dependent on modeled turbulence variables, is used to shift the closure from RANS to LES. Turbulent structures are initiated and sustained through the use of a recycling / rescaling technique. Two higher-order discretizations, the Piecewise Parabolic Method (PPM) of Colella and Woodward, and the SONIC-A ENO scheme of Suresh and Huyhn are used in the study. The results using the hybrid model show reasonably good agreement with time-averaged Mie scattering data and with experimental surface pressure distributions, even though the penetration of the jet into the cross-flow is slightly over-predicted. The LES/RANS results are used to examine the validity of commonly-used assumptions of constant Schmidt and Prandtl numbers in the intense mixing zone downstream of the injection location.

Boles, John A.↗
NASA NTRS2008

Initial Results from the Variable Intensity Sonic Boom Database

43 sonic booms generated (a few were evanescent waves) a) Overpressures of 0.08 to 2.20 lbf/sq ft; b) Rise-times of about 0.7 to 50 ms. Objectives: a) Structural response of a house of modern construction; b) Sonic boom propagation code validation. Approach: a) Measure shockwave directionality; b) Determine effect of height above ground on acoustic level; c) Generate atmospheric turbulence filter functions.

Haering, Edward A., Jr.↗
NASA NTRS2008

Output-Adaptive Tetrahedral Cut-Cell Validation for Sonic Boom Prediction

A cut-cell approach to Computational Fluid Dynamics (CFD) that utilizes the median dual of a tetrahedral background grid is described. The discrete adjoint is also calculated, which permits adaptation based on improving the calculation of a specified output (off-body pressure signature) in supersonic inviscid flow. These predicted signatures are compared to wind tunnel measurements on and off the configuration centerline 10 body lengths below the model to validate the method for sonic boom prediction. Accurate mid-field sonic boom pressure signatures are calculated with the Euler equations without the use of hybrid grid or signature propagation methods. Highly-refined, shock-aligned anisotropic grids were produced by this method from coarse isotropic grids created without prior knowledge of shock locations. A heuristic reconstruction limiter provided stable flow and adjoint solution schemes while producing similar signatures to Barth-Jespersen and Venkatakrishnan limiters. The use of cut-cells with an output-based adaptive scheme completely automated this accurate prediction capability after a triangular mesh is generated for the cut surface. This automation drastically reduces the manual intervention required by existing methods.

Park, Michael A.↗
NASA NTRS2011

Equivalent Longitudinal Area Distributions of the B-58 and XB-70-1 Airplanes for Use in Wave Drag and Sonic Boom Calculations

A detailed geometric description, in wave drag format, has been developed for the Convair B-58 and North American XB-70-1 delta wing airplanes. These descriptions have been placed on electronic files, the contents of which are described in this paper They are intended for use in wave drag and sonic boom calculations. Included in the electronic file and in the present paper are photographs and 3-view drawings of the two airplanes, tabulated geometric descriptions of each vehicle and its components, and comparisons of the electronic file outputs with existing data. The comparisons include a pictorial of the two airplanes based on the present geometric descriptions, and cross-sectional area distributions for both the normal Mach cuts and oblique Mach cuts above and below the vehicles. Good correlation exists between the area distributions generated in the late 1950s and 1960s and the present files. The availability of these electronic files facilitates further validation of sonic boom prediction codes through the use of two existing data bases on these airplanes, which were acquired in the 1960s and have not been fully exploited.

Tinetti, Ana F.↗
NASA NTRS2011

Measured Sonic Boom Signatures Above and Below the XB-70 Airplane Flying at Mach 1.5 and 37,000 Feet

During the 1966-67 Edwards Air Force Base (EAFB) National Sonic Boom Evaluation Program, a series of in-flight flow-field measurements were made above and below the USAF XB-70 using an instrumented NASA F-104 aircraft with a specially designed nose probe. These were accomplished in the three XB-70 flights at about Mach 1.5 at about 37,000 ft. and gross weights of about 350,000 lbs. Six supersonic passes with the F-104 probe aircraft were made through the XB-70 shock flow-field; one above and five below the XB-70. Separation distances ranged from about 3000 ft. above and 7000 ft. to the side of the XB-70 and about 2000 ft. and 5000 ft. below the XB-70. Complex near-field "sawtooth-type" signatures were observed in all cases. At ground level, the XB-70 shock waves had not coalesced into the two-shock classical sonic boom N-wave signature, but contained three shocks. Included in this report is a description of the generating and probe airplanes, the in-flight and ground pressure measuring instrumentation, the flight test procedure and aircraft positioning, surface and upper air weather observations, and the six in-flight pressure signatures from the three flights.

Maglieri, Domenic J.↗
NASA NTRS2011

Exhaust Nozzle Plume Effects on Sonic Boom Test Results for Isolated Nozzles

Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions were due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Recent work has been performed to reduce the magnitude of the sonic boom N-wave generated by airplane components with focus on shock waves caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analysis showed how the shock wave formed at the nozzle lip interacts with the nozzle boat-tail expansion wave. An experiment was conducted in the 1- by 1-ft Supersonic Wind Tunnel at the NASA Glenn Research Center to validate the computational study. Results demonstrated how the nozzle lip shock moved with increasing nozzle pressure ratio (NPR) and reduced the nozzle boat-tail expansion, causing a favorable change in the observed pressure signature. Experimental results were presented for comparison to the CFD results. The strong nozzle lip shock at high values of NPR intersected the nozzle boat-tail expansion and suppressed the expansion wave. Based on these results, it may be feasible to reduce the boat-tail expansion for a future supersonic aircraft with under-expanded nozzle exhaust flow by modifying nozzle pressure or nozzle divergent section geometry.

Castner, Raymond S.↗
NASA NTRS2012

Design of Rail Instrumentation for Wind Tunnel Sonic Boom Measurements and Computational-Experimental Comparisons

An innovative pressure rail concept for wind tunnel sonic boom testing of modern aircraft configurations with very low overpressures was designed with an adjoint-based solution-adapted Cartesian grid method. The computational method requires accurate free-air calculations of a test article as well as solutions modeling the influence of rail and tunnel walls. Specialized grids for accurate Euler and Navier-Stokes sonic boom computations were used on several test articles including complete aircraft models with flow-through nacelles. The computed pressure signatures are compared with recent results from the NASA 9- x 7-foot Supersonic Wind Tunnel using the advanced rail design.

Cliff, Susan E.↗
NASA NTRS2013

Using CFD Surface Solutions to Shape Sonic Boom Signatures Propagated from Off-Body Pressure

The conceptual design of a low-boom and low-drag supersonic aircraft remains a challenge despite significant progress in recent years. Inverse design using reversed equivalent area and adjoint methods have been demonstrated to be effective in shaping the ground signature propagated from computational fluid dynamics (CFD) off-body pressure distributions. However, there is still a need to reduce the computational cost in the early stages of design to obtain a baseline that is feasible for low-boom shaping, and in the search for a robust low-boom design over the entire sonic boom footprint. The proposed design method addresses the need to reduce the computational cost for robust low-boom design by using surface pressure distributions from CFD solutions to shape sonic boom ground signatures propagated from CFD off-body pressure.

Ordaz, Irian↗
NASA NTRS2014

Waveforms and Sonic Boom Perception and Response (WSPR): Low-Boom Community Response Program Pilot Test Design, Execution, and Analysis

The Waveforms and Sonic boom Perception and Response (WSPR) Program was designed to test and demonstrate the applicability and effectiveness of techniques to gather data relating human subjective response to multiple low-amplitude sonic booms. It was in essence a practice session for future wider scale testing on naive communities, using a purpose built low-boom demonstrator aircraft. The low-boom community response pilot experiment was conducted in California in November 2011. The WSPR team acquired sufficient data to assess and evaluate the effectiveness of the various physical and psychological data gathering techniques and analysis methods.

Page, Juliet A.↗
NASA NTRS2014

Specialized CFD Grid Generation Methods for Near-Field Sonic Boom Prediction

Ongoing interest in analysis and design of low sonic boom supersonic transports re- quires accurate and ecient Computational Fluid Dynamics (CFD) tools. Specialized grid generation techniques are employed to predict near- eld acoustic signatures of these con- gurations. A fundamental examination of grid properties is performed including grid alignment with ow characteristics and element type. The issues a ecting the robustness of cylindrical surface extrusion are illustrated. This study will compare three methods in the extrusion family of grid generation methods that produce grids aligned with the freestream Mach angle. These methods are applied to con gurations from the First AIAA Sonic Boom Prediction Workshop.

Park, Michael A.↗
NASA NTRS2015

Plume and Shock Interaction Effects on Sonic Boom in the 1-foot by 1-foot Supersonic Wind Tunnel

The desire to reduce or eliminate the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. Restrictions are due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed by the aircraft. A study has been performed focused on reducing the magnitude of the sonic boom N-wave generated by airplane components with a focus on shock waves caused by the exhaust nozzle plume. Testing was completed in the 1-foot by 1-foot supersonic wind tunnel to study the effects of an exhaust nozzle plume and shock wave interaction. The plume and shock interaction study was developed to collect data for computational fluid dynamics (CFD) validation of a nozzle plume passing through the shock generated from the wing or tail of a supersonic vehicle. The wing or tail was simulated with a wedgeshaped shock generator. This test entry was the first of two phases to collect schlieren images and off-body static pressure profiles. Three wedge configurations were tested consisting of strut-mounted wedges of 2.5- degrees and 5-degrees. Three propulsion configurations were tested simulating the propulsion pod and aft deck from a low boom vehicle concept, which also provided a trailing edge shock and plume interaction. Findings include how the interaction of the jet plume caused a thickening of the shock generated by the wedge (or aft deck) and demonstrate how the shock location moved with increasing nozzle pressure ratio.

Shock wave interaction↗
NASA NTRS2015

Unstructured Grids for Sonic Boom Analysis and Design

An evaluation of two methods for improving the process for generating unstructured CFD grids for sonic boom analysis and design has been conducted. The process involves two steps: the generation of an inner core grid using a conventional unstructured grid generator such as VGRID, followed by the extrusion of a sheared and stretched collar grid through the outer boundary of the core grid. The first method evaluated, known as COB, automatically creates a cylindrical outer boundary definition for use in VGRID that makes the extrusion process more robust. The second method, BG, generates the collar grid by extrusion in a very efficient manner. Parametric studies have been carried out and new options evaluated for each of these codes with the goal of establishing guidelines for best practices for maintaining boom signature accuracy with as small a grid as possible. In addition, a preliminary investigation examining the use of the CDISC design method for reducing sonic boom utilizing these grids was conducted, with initial results confirming the feasibility of a new remote design approach.

Campbell, Richard L.↗
NASA NTRS2015

Sonic Boom Pressure Signature Uncertainty Calculation and Propagation to Ground Noise

The objective of this study was to outline an approach for the quantification of uncertainty in sonic boom measurements and to investigate the effect of various near-field uncertainty representation approaches on ground noise predictions. These approaches included a symmetric versus asymmetric uncertainty band representation and a dispersion technique based on a partial sum Fourier series that allows for the inclusion of random error sources in the uncertainty. The near-field uncertainty was propagated to the ground level, along with additional uncertainty in the propagation modeling. Estimates of perceived loudness were obtained for the various types of uncertainty representation in the near-field. Analyses were performed on three configurations of interest to the sonic boom community: the SEEB-ALR, the 69o DeltaWing, and the LM 1021-01. Results showed that representation of the near-field uncertainty plays a key role in ground noise predictions. Using a Fourier series based dispersion approach can double the amount of uncertainty in the ground noise compared to a pure bias representation. Compared to previous computational fluid dynamics results, uncertainty in ground noise predictions were greater when considering the near-field experimental uncertainty.

West, Thomas K., IV↗