Results for “SUPERSONIC FLIGHT” — Page 14

NASA NTRS2025

Optimal Mission Profile for Low-Boom Supersonic Aircraft

An optimal mission profile, called the low-boom mission profile, is proposed for low-boom supersonic aircraft. The low-boom mission profile uses a fixed AoA for the cruise segment until reaching the cruise ceiling. The fixed AoA is determined by the cruise condition for the low-boom shaping of the aircraft. The derived CL trajectory is not necessarily the optimal trajectory for the cruise range when unrestricted supersonic flight is allowed. The undertrack sonic boom ground noise level is reduced during the cruise because of the increasing altitude from the fixed AoA trajectory and the approximate invariance of nondimensional overpressure at 3BL below a supersonic aircraft with respect to cruise altitude change when the AoA and cruise Mach are fixed. The approximate invariance of nondimensional overpressure at 3BL below a supersonic aircraft is verified using both Euler and Reynolds-averaged Navier-Stokes CFD solvers. A low-boom supersonic transport concept is used to illustrate that, compared to the fixed cruise altitude trajectory, the low-boom mission profile can increase the cruise range of a low-boom aircraft and reduce the undertrack sonic boom ground noise levels for the entire cruise segment. One significant contribution of this paper is the implication that an aircraft OML satisfying a sonic boom ground noise level constraint at SOC will satisfy the same noise level constraint over the entire cruise segment, as long as a fixed AoA trajectory is flown.

Low-boom characteristics↗

NASA NTRS1994

The effect of aircraft speed on the penetration of sonic boom noise into a flat ocean

As U.S. aircraft manufacturers now have focused their HSCT efforts on overwater supersonic flight, a great deal more must be known about sonic booms propagating overwater and interacting with the ocean. For example, it is thought that atmospheric turbulence effects are often much less severe over water than over land. Another important aspect of the overwater flight problems is the penetration of the sonic boom noise into the ocean, where there could be an environmental impact on sea life. This talk will present a brief review on the penetration of sonic boom noise into a large body of water with a flat surface. It has been determined recently that faster supersonic speeds imply greater penetration of sonic boom noise into the ocean. The new theory is derived from the original Sawyers paper and from the knowledge that for level flight a boom's duration is proportional to the quantity M/(M(exp 2)-1)(exp 3/8) where M is the Mach number. It is found that for depths of 10 m or less, the peak SPL varies less than 6 dB over a wide range of M. For greater depths, 100 m for example, increased Mach numbers may increase the SPL by 15 dB or more.

Sparrow, Victor W.↗

NASA NTRS1984

Flight investigation of various control inputs intended for parameter estimation

An experiment assessing the stability and control derivatives resulting from various control inputs was undertaken using the F-8 digital fly by wire aircraft. Improved control inputs have been proposed as a means of making stability and contol derivative estimation more efficient, thus reducing the cost of flight testing and data analysis. The subject inputs were either generated by the pilot or preprogrammed in a remote ground computer and telemetered to the aircraft. Nine preprogrammed inputs and three pilot generated inputs were assessed at subsonic and supersonic flight conditions, and both unaugmented and highly augmented flight control systems were used. Effects of input amplitude were also assessed. The inputs were divided into two general types - sinusoidal or with corners (a rapid and distinct change in slope). The inputs with corners, performed in the unaugmented mode, produced the best sets of stability and control derivatives. The simplest of these inputs, the pilot generated doublet, produced sets of derivatives as good as those produced by the more complex inputs. Small inputs produced worse derivatives than the unaugmented mode, and sinusoidal inputs produced worse derivatives than corner containing inputs.

Shafer, M. F.↗

NASA NTRS1951

Formulas for the Supersonic Loading, Lift, and Drag of Flat Swept-Back Wings with Leading Edges Behind the Mach Line

The method of superposition of linearized conical flows has been applied to the calculation of the aerodynamic properties, in supersonic flight, of thin flat, swept-back wings at an angle of attack. The wings are assumed to have rectilinear plan forms, with tips parallel to the stream, and to taper in the conventional sense. The investigation covers the moderately supersonic speed range where the Mach lines from the leading-edge apex lie ahead of the wing. The trailing edge may lie ahead of or behind the Mach lines from its apex. The case in which the Mach cone from one tip intersects the other tip is not treated. Formulas are obtained for the load distribution, the total lift, and the drag due to lift. For the cases in which the trailing edge is outside the Mach cone from its apex the formulas are complete. For wings with both leading and trailing edges behind their respective Mach lines, a degree of approximation is necessary. Charts of some of the functions derived are included to facilitate computing, and several examples are worked out in outline.

Cohen, Doris↗

NASA NTRS1987

Application of advanced computational codes in the design of an experiment for a supersonic throughflow fan rotor

Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The design of the experiment using advanced computational codes to calculate the components required is described. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier-Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.

Wood, Jerry R.↗

NASA NTRS1987

Application of advanced computational codes in the design of an experiment for a supersonic throughflow fan rotor

Increased emphasis on sustained supersonic or hypersonic cruise has revived interest in the supersonic throughflow fan as a possible component in advanced propulsion systems. Use of a fan that can operate with a supersonic inlet axial Mach number is attractive from the standpoint of reducing the inlet losses incurred in diffusing the flow from a supersonic flight Mach number to a subsonic one at the fan face. The design of the experiment using advanced computational codes to calculate the components required is described. The rotor was designed using existing turbomachinery design and analysis codes modified to handle fully supersonic axial flow through the rotor. A two-dimensional axisymmetric throughflow design code plus a blade element code were used to generate fan rotor velocity diagrams and blade shapes. A quasi-three-dimensional, thin shear layer Navier-Stokes code was used to assess the performance of the fan rotor blade shapes. The final design was stacked and checked for three-dimensional effects using a three-dimensional Euler code interactively coupled with a two-dimensional boundary layer code. The nozzle design in the expansion region was analyzed with a three-dimensional parabolized viscous code which corroborated the results from the Euler code. A translating supersonic diffuser was designed using these same codes.

Wood, Jerry R.↗

NASA NTRS1999

Exploring Flight Research with Experimental Gliders

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.

Source record↗

NASA NTRS2011

High-Quality Seismic Observations of Sonic Booms

The SonicBREWS project (Sonic Boom Resistant Earthquake Warning Systems) is a collaborative effort between Seismic Warning Systems, Inc. and NASA Dryden Flight Research Center. This project aims to evaluate the effects of sonic booms on Earthquake Warning Systems in order to prevent such systems from experiencing false alarms due to sonic booms. The airspace above the Antelope Valley, California includes the High Altitude Supersonic Corridor and the Black Mountain Supersonic Corridor. These corridors are among the few places in the US where supersonic flight is permitted, and sonic booms are commonplace in the Antelope Valley. One result of this project is a rich dataset of high-quality accelerometer records of sonic booms which can shed light on the interaction between these atmospheric phenomena and the solid earth. Nearly 100 sonic booms were recorded with low-noise triaxial MEMS accelerometers recording 1000 samples per second. The sonic booms had peak overpressures ranging up to approximately 10 psf and were recorded in three flight series in 2010 and 2011. Each boom was recorded with up to four accelerometers in various array configurations up to 100 meter baseline lengths, both in the built environment and the free field. All sonic booms were also recorded by nearby microphones. We present the results of the project in terms of the potential for sonic-boom-induced false alarms in Earthquake Warning Systems, and highlight some of the interesting features of the dataset.

Wurman, Gilead↗

NASA NTRS2020

Armstrong Flight Research Center Research, Technology & Engineering 2019

It is an honor to endorse the 2019 NASA’s Armstrong Flight Research Center Research, Technology, and Engineering Report. The talented researchers, engineers, and scientists at Armstrong continue to create innovative solutions to address some of the most pressing challenges facing the aerospace community, particularly as new technologies are expanding the field of flight transportation. Armstrong prides itself on its ability to develop new techniques, technologies, and test methodologies that address critical aerospace challenges. Researchers are currently undertaking cutting-edge projects that will advance our ability to achieve critical aerospace goals, such as quiet supersonic flight, electric aircraft, and integration of unmanned vehicles into the National Airspace System. In addition, teams are addressing some of the unique aspects associated with space access as well as other diverse challenges across mission directorates. As we continue our efforts to support NASA’s aerospace research and development missions, we strive to ensure that the public is aware of the advancements we have achieved. This report is a compilation of the wide range of work being conducted at Armstrong, along with contact information for the associated technologists responsible for each effort. We encourage you to reach out to these investigators for more information or to discuss collaboration ideas.

David F Voracek↗

NASA NTRS2008

A Grid Sourcing and Adaptation Study Using Unstructured Grids for Supersonic Boom Prediction

NASA created the Supersonics Project as part of the NASA Fundamental Aeronautics Program to advance technology that will make a supersonic flight over land viable. Computational flow solvers have lacked the ability to accurately predict sonic boom from the near to far field. The focus of this investigation was to establish gridding and adaptation techniques to predict near-to-mid-field (<10 body lengths below the aircraft) boom signatures at supersonic speeds using the USM3D unstructured grid flow solver. The study began by examining sources along the body the aircraft, far field sourcing and far field boundaries. The study then examined several techniques for grid adaptation. During the course of the study, volume sourcing was introduced as a new way to source grids using the grid generation code VGRID. Two different methods of using the volume sources were examined. The first method, based on manual insertion of the numerous volume sources, made great improvements in the prediction capability of USM3D for boom signatures. The second method (SSGRID), which uses an a priori adaptation approach to stretch and shear the original unstructured grid to align the grid and pressure waves, showed similar results with a more automated approach. Due to SSGRID s results and ease of use, the rest of the study focused on developing a best practice using SSGRID. The best practice created by this study for boom predictions using the CFD code USM3D involved: 1) creating a small cylindrical outer boundary either 1 or 2 body lengths in diameter (depending on how far below the aircraft the boom prediction is required), 2) using a single volume source under the aircraft, and 3) using SSGRID to stretch and shear the grid to the desired length.

Carter, Melissa B.↗

NASA NTRS1998

Shock Wave Dynamics in Weakly Ionized Gases

We have begun a comprehensive series of analyses and experiments to study the basic problem of shock wave dynamics in ionized media. Our objective is to isolate the mechanisms that are responsible for the decrease in the shock amplitude and also to determine the relevant plasma parameters that will be required for a drag reduction scheme in an actual high altitude hypersonic flight. Specifically, we have initiated a program of analyses and measurements with the objective of (i) fully characterizing the propagation dynamics in plasmas formed in gases of aerodynamic interest, (ii) isolating the mechanisms responsible for the decreased shock strength and increased shock velocity, (iii) extrapolating the laboratory observations to the technology of supersonic flight.

Johnson, Joseph A., III↗

NASA NTRS1993

Identification of integrated airframe-propulsion effects on an F-15 aircraft for application to drag minimization

The application of an adaptive real-time measurement-based performance optimization technique is being explored for a future flight research program. The key technical challenge of the approach is parameter identification, which uses a perturbation-search technique to identify changes in performance caused by forced oscillations of the controls. The controls on the NASA F-15 highly integrated digital electronic control (HIDEC) aircraft were perturbed using inlet cowl rotation steps at various subsonic and supersonic flight conditions to determine the effect on aircraft performance. The feasibility of the perturbation-search technique for identifying integrated airframe-propulsion system performance effects was successfully shown through flight experiments and postflight data analysis. Aircraft response and control data were analyzed postflight to identify gradients and to determine the minimum drag point. Changes in longitudinal acceleration as small as 0.004 g were measured, and absolute resolution was estimated to be 0.002 g or approximately 50 lbf of drag. Two techniques for identifying performance gradients were compared: a least-squares estimation algorithm and a modified maximum likelihood estimator algorithm. A complementary filter algorithm was used with the least squares estimator.

Schkolnik, Gerald S.↗

NASA NTRS1993

Identification of integrated airframe: Propulsion effects on an F-15 aircraft for application to drag minimization

The application of an adaptive real-time measurement-based performance optimization technique is being explored for a future flight research program. The key technical challenge of the approach is parameter identification, which uses a perturbation-search technique to identify changes in performance caused by forced oscillations of the controls. The controls on the NASA F-15 highly integrated digital electronic control (HIDEC) aircraft were perturbed using inlet cowl rotation steps at various subsonic and supersonic flight conditions to determine the effect on aircraft performance. The feasibility of the perturbation-search technique for identifying integrated airframe-propulsion system performance effects was successfully shown through flight experiments and postflight data analysis. Aircraft response and control data were analyzed postflight to identify gradients and to determine the minimum drag point. Changes in longitudinal acceleration as small as 0.004 g were measured, and absolute resolution was estimated to be 0.002 g or approximately 50 lbf of drag. Two techniques for identifying performance gradients were compared: a least-squares estimation algorithm and a modified maximum likelihood estimator algorithm. A complementary filter algorithm was used with the least squares estimator.

Schkolnik, Gerard S.↗

NASA NTRS2007

High-Heat-Flux Cyclic Durability of Thermal and Environmental Barrier Coatings

Advanced ceramic thermal and environmental barrier coatings will play an increasingly important role in future gas turbine engines because of their ability to protect the engine components and further raise engine temperatures. For the supersonic vehicles currently envisioned in the NASA fundamental aeronautics program, advanced gas turbine engines will be used to provide high power density thrust during the extended supersonic flight of the aircraft, while meeting stringent low emission requirements. Advanced ceramic coating systems are critical to the performance, life and durability of the hot-section components of the engine systems. In this work, the laser and burner rig based high-heat-flux testing approaches were developed to investigate the coating cyclic response and failure mechanisms under simulated supersonic long-duration cruise mission. The accelerated coating cracking and delamination mechanism under the engine high-heat-flux, and extended supersonic cruise time conditions will be addressed. A coating life prediction framework may be realized by examining the crack initiation and propagation in conjunction with environmental degradation under high-heat-flux test conditions.

Zhu, Dongming↗

NASA NTRS1995

Performance seeking control program overview

The Performance Seeking Control (PSC) program evolved from a series of integrated propulsion-flight control research programs flown at NASA Dryden Flight Research Center (DFRC) on an F-15. The first of these was the Digital Electronic Engine Control (DEEC) program and provided digital engine controls suitable for integration. The DEEC and digital electronic flight control system of the NASA F-15 were ideally suited for integrated controls research. The Advanced Engine Control System (ADECS) program proved that integrated engine and aircraft control could improve overall system performance. The objective of the PSC program was to advance the technology for a fully integrated propulsion flight control system. Whereas ADECS provided single variable control for an average engine, PSC controlled multiple propulsion system variables while adapting to the measured engine performance. PSC was developed as a model-based, adaptive control algorithm and included four optimization modes: minimum fuel flow at constant thrust, minimum turbine temperature at constant thrust, maximum thrust, and minimum thrust. Subsonic and supersonic flight testing were conducted at NASA Dryden covering the four PSC optimization modes and over the full throttle range. Flight testing of the PSC algorithm, conducted in a series of five flight test phases, has been concluded at NASA Dryden covering all four of the PSC optimization modes. Over a three year period and five flight test phases 72 research flights were conducted. The primary objective of flight testing was to exercise each PSC optimization mode and quantify the resulting performance improvements.

Orme, John S.↗

NASA NTRS2002

Replacement of Ablators with Phase-Change Material for Thermal Protection of STS Elements

As part of the research and development program to develop new Thermal Protection System (TPS) materials for aerospace applications at NASA's Marshall Space Flight Center (MSFC), an experimental study was conducted on a new concept for a non-ablative TPS material. Potential loss of TPS material and ablation by-products from the External Tank (ET) or Solid Rocket Booster (SRB) during Shuttle flight with the related Orbiter tile damage necessitates development of a non-ablative thermal protection system. The new Thermal Management Coating (TMC) consists of phase-change material encapsulated in micro spheres and a two-part resin system to adhere the coating to the structure material. The TMC uses a phase-change material to dissipate the heat produced during supersonic flight rather than an ablative material. This new material absorbs energy as it goes through a phase change during the heating portion of the flight profile and then the energy is slowly released as the phase-change material cools and returns to its solid state inside the micro spheres. The coating was subjected to different test conditions simulating design flight environments at the NASA/MSFC Improved Hot Gas Facility (IHGF) to study its performance.

Kaul, Raj K.↗

NASA NTRSDate not supplied

Jet Noise Flyover and Scale Model Tests

Renewed interest in commercial supersonic flight has rekindled the need for accurate jet-noise predictions as this source is believed to dominate at aircraft takeoff conditions. The current study compares scale-model data acquired in the NASA Aero-Acoustic Propulsion Laboratory with data obtained using a well-instrumented Learjet 25 in a flyover test completed in September 2022. The flight test included 73 flyovers with engine conditions ranging from 1.5 to 2.0 engine pressure ratios and flight Mach numbers between 0.24 and 0.27. Acoustic data were acquired with an 800-ft linear ground plate microphone array. Wind speed data were acquired up to 1000-ft altitude with a ground-based LiDAR system. Layered ambient temperature, pressure, and humidity were acquired with a weather drone. A 6% increase in the physical scale factor for the scale-model data was found to reasonably align the peak frequencies of the scale-model and flight data and resulted in peak levels for the scale model being roughly 0.7 dB above those for the flight data at NPR = 1.56 and roughly 1 dB below those for the flight data at NPR = 1.91 in the peak jet-noise direction. Comparisons with the SAE ARP876 model were poor especially at emission angles greater than, or equal to, 110° and at high frequencies.

jet noise, supersonic transport↗

NASA NTRS1985

A constant altitude flight survey method for mapping atmospheric ambient pressures and systematic radar errors

The flight test technique described uses controlled survey runs to determine horizontal atmospheric pressure variations and systematic altitude errors that result from space positioning measurements. The survey data can be used not only for improved air data calibrations, but also for studies of atmospheric structure and space positioning accuracy performance. The examples presented cover a wide range of radar tracking conditions for both subsonic and supersonic flight to an altitude of 42,000 ft.

Larson, T. J.↗