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At least 415 records · Page 23

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NASA NTRS2019

Control and Simulation of a Deployable Entry Vehicle with Aerodynamic Control Surfaces

In this paper, we investigate the static stability of a deployable entry vehicle called the Lifting Nano-ADEPT and design a control system to follow bank angle, angle-of-attack, and sideslip guidance commands. The control design, based on linear quadratic regulator optimal techniques, utilizes aerodynamic control surfaces to track angle-of-attack, sideslip angle, and bank angle commands. We demonstrate, using a nonlinear simulation environment, that the controller is able to accurately track step commands that may come from a guidance algorithm.

Margolis, Benjamin W. L.↗

NASA NTRS1973

Maximum likelihood identification and optimal input design for identifying aircraft stability and control derivatives

A new method of extracting aircraft stability and control derivatives from flight test data is developed based on the maximum likelihood cirterion. It is shown that this new method is capable of processing data from both linear and nonlinear models, both with and without process noise and includes output error and equation error methods as special cases. The first application of this method to flight test data is reported for lateral maneuvers of the HL-10 and M2/F3 lifting bodies, including the extraction of stability and control derivatives in the presence of wind gusts. All the problems encountered in this identification study are discussed. Several different methods (including a priori weighting, parameter fixing and constrained parameter values) for dealing with identifiability and uniqueness problems are introduced and the results given. The method for the design of optimal inputs for identifying the parameters of linear dynamic systems is also given. The criterion used for the optimization is the sensitivity of the system output to the unknown parameters. Several simple examples are first given and then the results of an extensive stability and control dervative identification simulation for a C-8 aircraft are detailed.

Stepner, D. E.↗

NASA NTRS1992

A study of transonic aerodynamic analysis methods for use with a hypersonic aircraft synthesis code

A means of performing routine transonic lift, drag, and moment analyses on hypersonic all-body and wing-body configurations were studied. The analysis method is to be used in conjunction with the Hypersonic Vehicle Optimization Code (HAVOC). A review of existing techniques is presented, after which three methods, chosen to represent a spectrum of capabilities, are tested and the results are compared with experimental data. The three methods consist of a wave drag code, a full potential code, and a Navier-Stokes code. The wave drag code, representing the empirical approach, has very fast CPU times, but very limited and sporadic results. The full potential code provides results which compare favorably to the wind tunnel data, but with a dramatic increase in computational time. Even more extreme is the Navier-Stokes code, which provides the most favorable and complete results, but with a very large turnaround time. The full potential code, TRANAIR, is used for additional analyses, because of the superior results it can provide over empirical and semi-empirical methods, and because of its automated grid generation. TRANAIR analyses include an all body hypersonic cruise configuration and an oblique flying wing supersonic transport.

Sandlin, Doral R.↗

NASA NTRS1995

Sensitivity Analysis of Wing Aeroelastic Responses

Design for prevention of aeroelastic instability (that is, the critical speeds leading to aeroelastic instability lie outside the operating range) is an integral part of the wing design process. Availability of the sensitivity derivatives of the various critical speeds with respect to shape parameters of the wing could be very useful to a designer in the initial design phase, when several design changes are made and the shape of the final configuration is not yet frozen. These derivatives are also indispensable for a gradient-based optimization with aeroelastic constraints. In this study, flutter characteristic of a typical section in subsonic compressible flow is examined using a state-space unsteady aerodynamic representation. The sensitivity of the flutter speed of the typical section with respect to its mass and stiffness parameters, namely, mass ratio, static unbalance, radius of gyration, bending frequency, and torsional frequency is calculated analytically. A strip theory formulation is newly developed to represent the unsteady aerodynamic forces on a wing. This is coupled with an equivalent plate structural model and solved as an eigenvalue problem to determine the critical speed of the wing. Flutter analysis of the wing is also carried out using a lifting-surface subsonic kernel function aerodynamic theory (FAST) and an equivalent plate structural model. Finite element modeling of the wing is done using NASTRAN so that wing structures made of spars and ribs and top and bottom wing skins could be analyzed. The free vibration modes of the wing obtained from NASTRAN are input into FAST to compute the flutter speed. An equivalent plate model which incorporates first-order shear deformation theory is then examined so it can be used to model thick wings, where shear deformations are important. The sensitivity of natural frequencies to changes in shape parameters is obtained using ADIFOR. A simple optimization effort is made towards obtaining a minimum weight design of the wing, subject to flutter constraints, lift requirement constraints for level flight and side constraints on the planform parameters of the wing using the IMSL subroutine NCONG, which uses successive quadratic programming.

Issac, Jason Cherian↗

NASA NTRS1982

Aerodynamics of an airfoil with a jet issuing from its surface

A simple, two dimensional, incompressible and inviscid model for the problem posed by a two dimensional wing with a jet issuing from its lower surface is considered and a parametric analysis is carried out to observe how the aerodynamic characteristics depend on the different parameters. The mathematical problem constitutes a boundary value problem where the position of part of the boundary is not known a priori. A nonlinear optimization approach was used to solve the problem, and the analysis reveals interesting characteristics that may help to better understand the physics involved in more complex situations in connection with high lift systems.

Tavella, D. A.↗

NASA NTRS1990

Aerodynamic design of the Cal Poly Da Vinci Human-Powered Helicopter

This paper will discuss the methodology used in designing the rotor and drive propellers for the third generation Cal Poly Da Vinci Human-Powered Helicopter. The rotor was designed using a lifting surface, uniform inflow hover analysis code and the propeller was designed using a minimum induced-loss method. Construction, geometry, and operating considerations are discussed as they impact the designs. Optimization of the design performance is also explained. The propellers were tested in a wind tunnel and results are compared with theoretical data. Successful flight tests of the Da Vinci III are discussed.

Larwood, Scott↗

NASA NTRS1986

Drag-reduction characteristics of aft-swept wing tips

Results are presented of a study conducted to analyze the potential aerodynamic improvements which can be obtained with aft-swept wing tips at subsonic speeds. A review and summary of previous swept-tip applications and studies are given. Additionally, the through evolution optimized tail-fin shapes of fast-cruising aquatic animals are examined as well as the wing planform shapes of sea birds. These superior long-distance travellers both display lifting-surface planform shapes with increased leading-edge sweep towards the tip. These observations and analytical results obtained with a nonlinear surface-panel method demonstrate that substantial induced-drag improvements can be obtained through the adoption of planform shapes with highly-tapered aft-swept tips.

Van Dam, C. P.↗

NASA NTRS2019

Performance Optimization of Plate Airfoils for Martian Rotor Applications Using a Genetic Algorithm

The Mars Helicopter Technology Demonstrator will be flying on the NASA Mars 2020 rover mission scheduled to launch in July of 2020. The goal is to demonstrate the viability and potential of heavier-than-air vehicles in the Martian atmosphere. Research is performed at the Jet Propulsion Laboratory and NASA Ames Research Center to extend these capabilities and develop the Mars Science Helicopter as the next possible step for Martian rotorcraft. The Mars Science Helicopter mass is scaled up to the 5 to 20 kg range, allowing a greater payload (approximately 0.5 to 2.0 kg), and greater range (approximately 3 km). Key to achieving these targets is careful aerodynamic rotor design. The Martian atmosphere’s low density and the small helicopter rotors result in very low chord-based Reynolds number flows, which reduces rotor performance. A continuous genetic algorithm is developed to optimize airfoil shapes at representative conditions for the Martian atmosphere. Previous research indicates that sharp leading edges and plate-like airfoils can out-perform conventional airfoil shapes. The present optimization allows for camber and thickness variation of curved and polygonal thin airfoils with sharp leading edges. The airfoil performance is evaluated at the highest attainable liftto- drag ratio near a moderate lift coefficient at compressible Mach numbers, as expected for Martian rotor application. Increases between 16% and 29% in airfoil lift-to-drag ratio at fixed lift coefficients are observed when compared with the Mars Helicopter Technology Demonstrator airfoils. Improvements in hover figure of merit are estimated to be between 4% and 10%, when applied to the Mars Helicopter Technology Demonstrator.

Koning, Witold J. F.↗

NASA NTRS1984

Fuel conservative guidance concept for shipboard landing of powered-life aircraft

A simulation study was undertaken to investigate the application of energy conservative guidance (ECG) software, developed at NASA Ames Research Center, to improve the time and fuel efficiency of powered lift airplanes operating from aircraft carriers at sea. When a flightpath is indicated by a set of initial conditions for the aircraft and a set of positional waypoints with associated airspeeds, the ECG software synthesizes the necessary guidance commands to optimize fuel and time along the specified path. A major feature of the ECG system is the ability to synthesize a trajectory that will allow the aircraft to capture the specified path at any waypoint with the desired heading and airspeed from an arbitrary set of initial conditions. Five paths were identified and studied. These paths demonstrate the ECG system's ability to save flight time and fuel by more efficiently managing the aircraft's capabilities. Results of this simulation study show that when restrictions on the approach flightpath imposed for manual operation are removed completely, fuel consumption during the approach was reduced by as much as 49% (610 lb fuel) and the time required to fly the flightpath was reduced by as much as 41% (5 min). Savings due to ECG were produced by: (1) shortening the total flight time; (2) keeping the airspeed high as long as possible to minimize time spent flying in a regime in which more engine thrust is required for lift to aid the aerodynamic lift; (3) minimizing time spent flying at constant altitude at slow airspeeds; and (4) synthesizing a path from any location for a direct approach to landing without entering a holding pattern or other fixed approach path.

Warner, D. N., Jr.↗

NASA NTRS1979

Inverse boundary-layer technique for airfoil design

A description is presented of a technique for the optimization of airfoil pressure distributions using an interactive inverse boundary-layer program. This program allows the user to determine quickly a near-optimum subsonic pressure distribution which meets his requirements for lift, drag, and pitching moment at the desired flow conditions. The method employs an inverse turbulent boundary-layer scheme for definition of the turbulent recovery portion of the pressure distribution. Two levels of pressure-distribution architecture are used - a simple roof top for preliminary studies and a more complex four-region architecture for a more refined design. A technique is employed to avoid the specification of pressure distributions which result in unrealistic airfoils, that is, those with negative thickness. The program allows rapid evaluation of a designed pressure distribution off-design in Reynolds number, transition location, and angle of attack, and will compute an airfoil contour for the designed pressure distribution using linear theory.

Henderson, M. L.↗

NASA NTRS1987

Optimal trajectories for aeroassisted, noncoplanar orbital transfer. II - LEO-to-LEO transfer

Both classical and minimax problems of optimal control arising in the study of noncoplanar, aeroassisted orbital transfer are considered and are illustrated with the example of LEO-to-LEO transfer. Trajectory control is achieved by modulation of the lift coefficient and the angle of bank. Problems considered include the minimization of the energy required for orbital transfer, maximization of the flight time during the atmospheric portion of the trajectory, and minimization of the peak heating rate. The near-grazing solution is found to be a good compromise between energy and heating requirements.

Miele, A.↗

NASA NTRS1995

Numerical Methodology For Full-Body TPS Sizing And Optimization For Access-To-Space Vehicles

This paper presents details of the methodology and numerical procedures developed at NASA Ames for full-body TEXAS sizing and optimization for Access to Space vehicle concepts. The core of the procedures is a robust implicit solver for one dimensional transient heat conduction in reusable multilayer TEXAS stackups. The solver includes an arbitrary number of material layers, contact resistances between materials, temperature and pressure dependent material and surface properties, numerous boundary-condition options, and self-adaptive time stepping. The solver is coupled with the Access-to-Space material database of 23 candidate TPS and structural materials and a thermal-environment database obtained from trajectory-based fullbody Navier-Stokes computations of the external flowfield. The thermal environment and material response are coupled through the use of T-type heat transfer coefficients. TIPS sizing and weight optimization are performed at every surface point on the vehicle based on sizing constraints which include material temperature limits, maximum backwall temperature, and cumulative interior heat flux. Typical results are presented for a lifting body concept with 10000 surface points, which required 35 minutes to compute on an SGI Indigo 2.

Milos, Frank S.↗

NASA NTRS2010

Evaluation of Dual-Launch Lunar Architectures Using the Mission Assessment Post Processor

The National Aeronautics and Space Administrations (NASA) Constellation Program is currently designing a new transportation system to replace the Space Shuttle, support human missions to both the International Space Station (ISS) and the Moon, and enable the eventual establishment of an outpost on the lunar surface. The present Constellation architecture is designed to meet nominal capability requirements and provide flexibility sufficient for handling a host of contingency scenarios including (but not limited to) launch delays at the Earth. This report summarizes a body of work performed in support of the Review of U.S. Human Space Flight Committee. It analyzes three lunar orbit rendezvous dual-launch architecture options which incorporate differing methodologies for mitigating the effects of launch delays at the Earth. NASA employed the recently-developed Mission Assessment Post Processor (MAPP) tool to quickly evaluate vehicle performance requirements for several candidate approaches for conducting human missions to the Moon. The MAPP tool enabled analysis of Earth perturbation effects and Earth-Moon geometry effects on the integrated vehicle performance as it varies over the 18.6-year lunar nodal cycle. Results are provided summarizing best-case and worst-case vehicle propellant requirements for each architecture option. Additionally, the associated vehicle payload mass requirements at launch are compared between each architecture and against those of the Constellation Program. The current Constellation Program architecture assumes that the Altair lunar lander and Earth Departure Stage (EDS) vehicles are launched on a heavy lift launch vehicle. The Orion Crew Exploration Vehicle (CEV) is separately launched on a smaller man-rated vehicle. This strategy relaxes man-rating requirements for the heavy lift launch vehicle and has the potential to significantly reduce the cost of the overall architecture over the operational lifetime of the program. The crew launch occurs first, four days prior to the optimal trans-lunar injection (TLI) departure window. This is done to allow for launch delays in the Altair/EDS launch. During this time, the Orion vehicle is required to conduct orbit maintenance while loitering in low Earth orbit (LEO). The alternative architectures presented aim to eliminate the need for costly orbit maintenance maneuvers while loitering in LEO. In all of the alternative architectures considered, it is assumed that the Altair and Orion vehicles are nominally launched 90 minutes apart, depart the Earth separately, and complete the rendezvous and docking sequence at the Moon. In this lunar orbit rendezvous (LOR) strategy, both the Altair and Orion vehicles will require separate EDS stages, and each will be required to perform lunar orbit insertion (LOI). This has the effect of balancing payload requirements between the two launch vehicles at the Earth. In this case, the overall payload mass is increased slightly, but the increased mission costs could potentially be offset by requiring the construction of two rockets similar in size and nature, unlike the current Constellation architecture. Three dual-launch architecture options with LOR were evaluated, which incorporate differing methodologies for mitigating the effects of launch delays at the Earth. Benefits and drawbacks of each of the dual-launch architecture options with LOR are discussed and the overall mission performance is compared with that of the existing Constellation Program lunar architecture.

Stewart, Shaun M.↗

NASA NTRS1975

Methods Studies on System Identification from Transient Rotor Tests

Some of the more important methods are discussed that have been used or proposed for aircraft parameter identification. The methods are classified into two groups: Equation error or regression estimates and Bayesian estimates and their derivatives that are based on probabilistic concepts. In both of these two groups the cost function can be optimized either globally over the entire time span of the transient, or sequentially, leading to the formulation of optimum filters. Identifiability problems and the validation of the estimates are briefly outlined, and applications to lifting rotors are discussed.

Hohenemser, K. H.↗

NASA NTRSDate not supplied

TPSAS-NF1676L-30697-DND

The goal of the NASA X-57 Electric Airplane was to utilize the versatility and efficiency of electric propulsion to design a cruise optimized airplane. The wing planform was reduced 42%. A slotted flap and distributed fan system were developed to achieve CL?s in excess of 4.0. Performance objectives, features of the wing and high-lift system, as well as integration considerations will be discussed. CFD performance analysis and details of the integrated structure will be discussed.

Jeff Viken↗

NASA NTRSDate not supplied

Design of a Propeller with Global Minimum Torque

Academia, industry, and Government are actively working towards a future where dozens of small to large Unmanned Aerial Systems (UAS) are present within urban environments. In this environment, high noise pollution and inefficiencies have the potential to render these new technologies impractical due to public annoyance and nonacceptance. This study aims to design and analyze a novel propeller design for minimum global torque using an optimization algorithm to find the full three dimensional solution of maximum efficiency. This resulted in a 𝐶 𝑙 distribution defining the global minimum torque solution for propeller design which corresponded with moving as much lift inboard as possible and reducing lift rapidly at the blade tip. The novel propeller (dubbed "Prandtl" propeller) was then compared with the current gold standard in propeller design, a minimum induced loss (MIL) propeller, which had all the same geometric properties except for blade twist and produced the same amount of thrust at the same advance ratio. The results of multiple iterations of Prandtl propellers showed an increase in efficiency of 2.0-3.3% for the Prandtl blade when compared to the equivalent MIL blade. A potential added benefit of this blade design is lower noise generation due to the lower lift loading at the tip reducing the large shear layer intensity which is the point source of noise in propellers. Two major breakthroughs for enabling widespread use of UAS in urban environments are noise reduction and vehicle efficiency, and this new propeller design has the potential to provide both.

Propeller↗

NASA NTRSDate not supplied

Design of a Propeller with Global Minimum Torque

Propeller↗

NASA NTRS2002

Prediction of Aerodynamic Coefficient using Genetic Algorithm Optimized Neural Network for Sparse Data

Wind tunnels use scale models to characterize aerodynamic coefficients, Wind tunnel testing can be slow and costly due to high personnel overhead and intensive power utilization. Although manual curve fitting can be done, it is highly efficient to use a neural network to define the complex relationship between variables. Numerical simulation of complex vehicles on the wide range of conditions required for flight simulation requires static and dynamic data. Static data at low Mach numbers and angles of attack may be obtained with simpler Euler codes. Static data of stalled vehicles where zones of flow separation are usually present at higher angles of attack require Navier-Stokes simulations which are costly due to the large processing time required to attain convergence. Preliminary dynamic data may be obtained with simpler methods based on correlations and vortex methods; however, accurate prediction of the dynamic coefficients requires complex and costly numerical simulations. A reliable and fast method of predicting complex aerodynamic coefficients for flight simulation I'S presented using a neural network. The training data for the neural network are derived from numerical simulations and wind-tunnel experiments. The aerodynamic coefficients are modeled as functions of the flow characteristics and the control surfaces of the vehicle. The basic coefficients of lift, drag and pitching moment are expressed as functions of angles of attack and Mach number. The modeled and training aerodynamic coefficients show good agreement. This method shows excellent potential for rapid development of aerodynamic models for flight simulation. Genetic Algorithms (GA) are used to optimize a previously built Artificial Neural Network (ANN) that reliably predicts aerodynamic coefficients. Results indicate that the GA provided an efficient method of optimizing the ANN model to predict aerodynamic coefficients. The reliability of the ANN using the GA includes prediction of aerodynamic coefficients to an accuracy of 110% . In our problem, we would like to get an optimized neural network architecture and minimum data set. This has been accomplished within 500 training cycles of a neural network. After removing training pairs (outliers), the GA has produced much better results. The neural network constructed is a feed forward neural network with a back propagation learning mechanism. The main goal has been to free the network design process from constraints of human biases, and to discover better forms of neural network architectures. The automation of the network architecture search by genetic algorithms seems to have been the best way to achieve this goal.

Rajkumar, T.↗