Design Methods and Optimization for Morphing Aircraft
NASA Technical Reports Server (NTRS)
Crossley, William A.
2005-01-01
This report provides a summary of accomplishments made during this research effort. The major accomplishments are in three areas. The first is the use of a multiobjective optimization strategy to help identify potential morphing features that uses an existing aircraft sizing code to predict the weight, size and performance of several fixed-geometry aircraft that are Pareto-optimal based upon on two competing aircraft performance objectives. The second area has been titled morphing as an independent variable and formulates the sizing of a morphing aircraft as an optimization problem in which the amount of geometric morphing for various aircraft parameters are included as design variables. This second effort consumed most of the overall effort on the project. The third area involved a more detailed sizing study of a commercial transport aircraft that would incorporate a morphing wing to possibly enable transatlantic point-to-point passenger service.
Aircraft family design using enhanced collaborative optimization
NASA Astrophysics Data System (ADS)
Roth, Brian Douglas
Significant progress has been made toward the development of multidisciplinary design optimization (MDO) methods that are well-suited to practical large-scale design problems. However, opportunities exist for further progress. This thesis describes the development of enhanced collaborative optimization (ECO), a new decomposition-based MDO method. To support the development effort, the thesis offers a detailed comparison of two existing MDO methods: collaborative optimization (CO) and analytical target cascading (ATC). This aids in clarifying their function and capabilities, and it provides inspiration for the development of ECO. The ECO method offers several significant contributions. First, it enhances communication between disciplinary design teams while retaining the low-order coupling between them. Second, it provides disciplinary design teams with more authority over the design process. Third, it resolves several troubling computational inefficiencies that are associated with CO. As a result, ECO provides significant computational savings (relative to CO) for the test cases and practical design problems described in this thesis. New aircraft development projects seldom focus on a single set of mission requirements. Rather, a family of aircraft is designed, with each family member tailored to a different set of requirements. This thesis illustrates the application of decomposition-based MDO methods to aircraft family design. This represents a new application area, since MDO methods have traditionally been applied to multidisciplinary problems. ECO offers aircraft family design the same benefits that it affords to multidisciplinary design problems. Namely, it simplifies analysis integration, it provides a means to manage problem complexity, and it enables concurrent design of all family members. In support of aircraft family design, this thesis introduces a new wing structural model with sufficient fidelity to capture the tradeoffs associated with component
Application of heuristic optimization in aircraft design
NASA Astrophysics Data System (ADS)
Hu, Zhenning
Genetic algorithms and the related heuristic optimization strategies are introduced and their applications in the aircraft design are developed. Generally speaking, genetic algorithms belong to non-deterministic direct search methods, which are most powerful in finding optimum or near-optimum solutions of a very complex system where a little priori knowledge is known. Therefore they have a wide application in aerospace systems. Two major aircraft optimal design projects are illustrated in this dissertation. The first is the application of material optimization of aligned fiber laminate composites in the presence of stress concentrations. After a large number of tests on laminates with different layers, genetic algorithms find an alignment pattern in a certain range for the Boeing Co. specified material. The second project is the application of piezoelectric actuator placement on a generic tail skins to reduce the 2nd mode vibration caused by buffet, which is part of a Boeing project to control the buffet effect on aircraft. In this project, genetic algorithms are closely involved with vibration analysis and finite element analysis. Actuator optimization strategies are first tested on the theoretical beam models to gain experience, and then the generic tail model is applied. Genetic algorithms achieve a great success in optimizing up to 888 actuator parameters on the tail skins.
A robust optimization methodology for preliminary aircraft design
NASA Astrophysics Data System (ADS)
Prigent, S.; MarĂ©chal, P.; Rondepierre, A.; Druot, T.; Belleville, M.
2016-05-01
This article focuses on a robust optimization of an aircraft preliminary design under operational constraints. According to engineers' know-how, the aircraft preliminary design problem can be modelled as an uncertain optimization problem whose objective (the cost or the fuel consumption) is almost affine, and whose constraints are convex. It is shown that this uncertain optimization problem can be approximated in a conservative manner by an uncertain linear optimization program, which enables the use of the techniques of robust linear programming of Ben-Tal, El Ghaoui, and Nemirovski [Robust Optimization, Princeton University Press, 2009]. This methodology is then applied to two real cases of aircraft design and numerical results are presented.
Multidisciplinary Optimization Methods for Aircraft Preliminary Design
NASA Technical Reports Server (NTRS)
Kroo, Ilan; Altus, Steve; Braun, Robert; Gage, Peter; Sobieski, Ian
1994-01-01
This paper describes a research program aimed at improved methods for multidisciplinary design and optimization of large-scale aeronautical systems. The research involves new approaches to system decomposition, interdisciplinary communication, and methods of exploiting coarse-grained parallelism for analysis and optimization. A new architecture, that involves a tight coupling between optimization and analysis, is intended to improve efficiency while simplifying the structure of multidisciplinary, computation-intensive design problems involving many analysis disciplines and perhaps hundreds of design variables. Work in two areas is described here: system decomposition using compatibility constraints to simplify the analysis structure and take advantage of coarse-grained parallelism; and collaborative optimization, a decomposition of the optimization process to permit parallel design and to simplify interdisciplinary communication requirements.
A multiple objective optimization approach to aircraft control systems design
NASA Technical Reports Server (NTRS)
Tabak, D.; Schy, A. A.; Johnson, K. G.; Giesy, D. P.
1979-01-01
The design of an aircraft lateral control system, subject to several performance criteria and constraints, is considered. While in the previous studies of the same model a single criterion optimization, with other performance requirements expressed as constraints, has been pursued, the current approach involves a multiple criteria optimization. In particular, a Pareto optimal solution is sought.
Formulation of the multimission aircraft design optimization problem
NASA Astrophysics Data System (ADS)
Straussfogel, Dennis M.
1998-12-01
The conventional single-mission aircraft design optimization problem is reformulated to allow design and optimization for multiple missions. Defining the aircraft mission as a set of continuous scalar variables leads to the concepts of mission vector and mission space. The multimission aircraft design optimization problem becomes one of optimizing a design for several different points in the mission space, simultaneously. In the limit, a design can be optimized simultaneously for all points in the mission space. Mapping various points from the optimization design space into the mission space generates actual and theoretical optimum performance surfaces. The multimission optimum configuration is defined as the single configuration that minimizes the difference between the actual performance and the theoretical optimum performance. The multimission aircraft design optimization method can be applied over several distinct mission by summing the differences between actual and theoretical optimum performance, or over all possible missions by integrating the difference between the actual and theoretical optimum performance surfaces. The concepts associated with the mission vector, mission space, and multimission optimum configuration are presented in mathematical form. The objective function for the multimission optimization problem is expressed both as a summation over discrete mission points and as an integral over the entire mission space. Mathematical expressions for objective functions based on both single and multiple objectives are developed and presented. A weighting function, emphasizing certain parts of the mission space over others, is also discussed. The multimission aircraft design optimization method is applied to an elementary wing-design optimization problem for an executive jet. The optimization problem is solved numerically for single and multiple objectives, with and without a functional constraint. The effect of the different objective functions and
Multidisciplinary design optimization of low-noise transport aircraft
NASA Astrophysics Data System (ADS)
Leifsson, Leifur Thor
The objective of this research is to examine how to design low-noise transport aircraft using Multidisciplinary Design Optimization (MDO). The subject is approached by designing for low-noise both implicitly and explicitly. The explicit design approach involves optimizing an aircraft while explicitly constraining the noise level. An MDO framework capable of optimizing both a cantilever wing and a Strut-Braced-Wing (SBW) aircraft was developed. The objective is to design aircraft for low-airframe-noise at the approach conditions and quantify the change in weight and performance with respect to a traditionally designed aircraft. The results show that reducing airframe noise by reducing approach speed alone, will not provide significant noise reduction without a large performance and weight penalty. Therefore, more dramatic changes to the aircraft design are needed to achieve a significant airframe noise reduction. Another study showed that the trailing-edge flap can be eliminated, as well as all the noise associated with that device, without incurring a significant weight and performance penalty. Lastly, an airframe noise analysis showed that a SBW aircraft with short fuselage-mounted landing gear could have a similar or potentially a lower airframe noise level than a comparable cantilever wing aircraft. The implicit design approach involves selecting a configuration that supports a low-noise operation, and optimizing for performance. In this study a Blended-Wing-Body (BWB) transport aircraft, with a conventional and a distributed propulsion system, was optimized for minimum take-off gross weight. The effects of distributed propulsion were studied using an MDO framework previously developed at Virginia Tech. The results show that more than two thirds of the theoretical savings of distributed propulsion are required for the BWB designs with a distributed propulsion system to have comparable gross weight as those with a conventional propulsion system. Therefore
Experiences performing conceptual design optimization of transport aircraft
NASA Technical Reports Server (NTRS)
Arbuckle, P. D.; Sliwa, S. M.
1984-01-01
Optimum Preliminary Design of Transports (OPDOT) is a computer program developed at NASA Langley Research Center for evaluating the impact of new technologies upon transport aircraft. For example, it provides the capability to look at configurations which have been resized to take advantage of active controls and provide and indication of economic sensitivity to its use. Although this tool returns a conceptual design configuration as its output, it does not have the accuracy, in absolute terms, to yield satisfactory point designs for immediate use by aircraft manufacturers. However, the relative accuracy of comparing OPDOT-generated configurations while varying technological assumptions has been demonstrated to be highly reliable. Hence, OPDOT is a useful tool for ascertaining the synergistic benefits of active controls, composite structures, improved engine efficiencies and other advanced technological developments. The approach used by OPDOT is a direct numerical optimization of an economic performance index. A set of independent design variables is iterated, given a set of design constants and data. The design variables include wing geometry, tail geometry, fuselage size, and engine size. This iteration continues until the optimum performance index is found which satisfies all the constraint functions. The analyst interacts with OPDOT by varying the input parameters to either the constraint functions or the design constants. Note that the optimization of aircraft geometry parameters is equivalent to finding the ideal aircraft size, but with more degrees of freedom than classical design procedures will allow.
Optimizing conceptual aircraft designs for minimum life cycle cost
NASA Technical Reports Server (NTRS)
Johnson, Vicki S.
1989-01-01
A life cycle cost (LCC) module has been added to the FLight Optimization System (FLOPS), allowing the additional optimization variables of life cycle cost, direct operating cost, and acquisition cost. Extensive use of the methodology on short-, medium-, and medium-to-long range aircraft has demonstrated that the system works well. Results from the study show that optimization parameter has a definite effect on the aircraft, and that optimizing an aircraft for minimum LCC results in a different airplane than when optimizing for minimum take-off gross weight (TOGW), fuel burned, direct operation cost (DOC), or acquisition cost. Additionally, the economic assumptions can have a strong impact on the configurations optimized for minimum LCC or DOC. Also, results show that advanced technology can be worthwhile, even if it results in higher manufacturing and operating costs. Examining the number of engines a configuration should have demonstrated a real payoff of including life cycle cost in the conceptual design process: the minimum TOGW of fuel aircraft did not always have the lowest life cycle cost when considering the number of engines.
Design optimization of high-speed proprotor aircraft
NASA Technical Reports Server (NTRS)
Schleicher, David R.; Phillips, James D.; Carbajal, Kevin B.
1993-01-01
NASA's high-speed rotorcraft (HSRC) studies have the objective of investigating technology for vehicles that have both low downwash velocities and forward flight speed capability of up to 450 knots. This paper investigates a tilt rotor, a tilt wing, and a folding tilt rotor designed for a civil transport mission. Baseline aircraft models using current technology are developed for each configuration using a vertical/short takeoff and landing (V/STOL) aircraft design synthesis computer program to generate converged vehicle designs. Sensitivity studies and numerical optimization are used to illustrate each configuration's key design tradeoffs and constraints. Minimization of the gross takeoff weight is used as the optimization objective function. Several advanced technologies are chosen, and their relative impact on future configurational development is discussed. Finally, the impact of maximum cruise speed on vehicle figures of merit (gross weight, productivity, and direct operating cost) is analyzed. The three most important conclusions from the study are payload ratios for these aircraft will be commensurate with current fixed-wing commuter aircraft; future tilt rotors and tilt wings will be significantly lighter, more productive, and cheaper than competing folding tilt rotors; and the most promising technologies are an advanced-technology proprotor for both tilt rotor and tilt wing and advanced structural materials for the folding tilt rotor.
Design-oriented aeroservoelastic optimization of strain-actuated aircraft
NASA Astrophysics Data System (ADS)
Jackson, Timothy W.
An integrated design-oriented aeroservoelastic optimization capability for strain-actuated aircraft is presented. This capability is called SMART and it encompasses a suite of computer applications created for conceptual and preliminary design of aircraft augmented with "smart" actuation technologies. The SMART suite of applications includes: (1) a dedicated pre-processor for vehicle geometry, material, actuator, mechanism, and sensor layout; (2) a dedicated finite element automesher for conventional and strain-actuated flight vehicles; (3) integration of structural dynamics with a state of the art commercial unsteady aerodynamics code (ZAERO) via automated pre- and post-processors; (4) a database architecture for analyzing multiple designs and flight conditions; and (5) automated open- and closed-loop aeroservoelastic (ASE) model preparation. The analysis techniques used as the basis for SMART are suitable (within the range of application of linear theory) for modeling real flight vehicles with real large-scale structural, aerodynamic, and control systems. These techniques include: (6) dedicated linear finite element infrastructure for modeling conventional and strain-actuated (temperature and voltage induced) flight structures; (7) dedicated static and dynamic finite element solvers; (8) state space stability analysis for coupled aeroservoelastic systems; (9) computational tools for LQR controller design; and (10) analysis tools for the calculation of random response of linear systems to random inputs. SMART can compute: (11) static aeroelastic deformations and stresses in trimmed maneuvering elastic airplanes; (12) open-loop aeroelastic poles; (13) closed-loop (control by LQR) aeroservoelastic poles; (14) gust response (random gusts) of the open-loop and closed-loop aeroservoelastic system, and, also, when aerodynamic loads are not present; and (15) deformation and stresses of passive or actively-controlled structures subject to dynamic and static loads. To
Aircraft design for mission performance using nonlinear multiobjective optimization methods
NASA Technical Reports Server (NTRS)
Dovi, Augustine R.; Wrenn, Gregory A.
1990-01-01
A new technique which converts a constrained optimization problem to an unconstrained one where conflicting figures of merit may be simultaneously considered was combined with a complex mission analysis system. The method is compared with existing single and multiobjective optimization methods. A primary benefit from this new method for multiobjective optimization is the elimination of separate optimizations for each objective, which is required by some optimization methods. A typical wide body transport aircraft is used for the comparative studies.
NASA Technical Reports Server (NTRS)
Bowers, Albion H. (Inventor); Uden, Edward (Inventor)
2016-01-01
The present invention is an aircraft wing design that creates a bell shaped span load, which results in a negative induced drag (induced thrust) on the outer portion of the wing; such a design obviates the need for rudder control of an aircraft.
NASA Technical Reports Server (NTRS)
Rommel, Bruce A.
1989-01-01
An overview of the Aeroelastic Design Optimization Program (ADOP) at the Douglas Aircraft Company is given. A pilot test program involving the animation of mode shapes with solid rendering as well as wire frame displays, a complete aircraft model of a high-altitude hypersonic aircraft to test ADOP procedures, a flap model, and an aero-mesh modeler for doublet lattice aerodynamics are discussed.
Optimization applications in aircraft engine design and test
NASA Technical Reports Server (NTRS)
Pratt, T. K.
1984-01-01
Starting with the NASA-sponsored STAEBL program, optimization methods based primarily upon the versatile program COPES/CONMIN were introduced over the past few years to a broad spectrum of engineering problems in structural optimization, engine design, engine test, and more recently, manufacturing processes. By automating design and testing processes, many repetitive and costly trade-off studies have been replaced by optimization procedures. Rather than taking engineers and designers out of the loop, optimization has, in fact, put them more in control by providing sophisticated search techniques. The ultimate decision whether to accept or reject an optimal feasible design still rests with the analyst. Feedback obtained from this decision process has been invaluable since it can be incorporated into the optimization procedure to make it more intelligent. On several occasions, optimization procedures have produced novel designs, such as the nonsymmetric placement of rotor case stiffener rings, not anticipated by engineering designers. In another case, a particularly difficult resonance contraint could not be satisfied using hand iterations for a compressor blade, when the STAEBL program was applied to the problem, a feasible solution was obtained in just two iterations.
Optimal Input Design for Aircraft Parameter Estimation using Dynamic Programming Principles
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.; Klein, Vladislav
1990-01-01
A new technique was developed for designing optimal flight test inputs for aircraft parameter estimation experiments. The principles of dynamic programming were used for the design in the time domain. This approach made it possible to include realistic practical constraints on the input and output variables. A description of the new approach is presented, followed by an example for a multiple input linear model describing the lateral dynamics of a fighter aircraft. The optimal input designs produced by the new technique demonstrated improved quality and expanded capability relative to the conventional multiple input design method.
Optimal input design for aircraft parameter estimation using dynamic programming principles
NASA Technical Reports Server (NTRS)
Klein, Vladislav; Morelli, Eugene A.
1990-01-01
A new technique was developed for designing optimal flight test inputs for aircraft parameter estimation experiments. The principles of dynamic programming were used for the design in the time domain. This approach made it possible to include realistic practical constraints on the input and output variables. A description of the new approach is presented, followed by an example for a multiple input linear model describing the lateral dynamics of a fighter aircraft. The optimal input designs produced by the new technique demonstrated improved quality and expanded capability relative to the conventional multiple input design method.
A Subsonic Aircraft Design Optimization With Neural Network and Regression Approximators
NASA Technical Reports Server (NTRS)
Patnaik, Surya N.; Coroneos, Rula M.; Guptill, James D.; Hopkins, Dale A.; Haller, William J.
2004-01-01
The Flight-Optimization-System (FLOPS) code encountered difficulty in analyzing a subsonic aircraft. The limitation made the design optimization problematic. The deficiencies have been alleviated through use of neural network and regression approximations. The insight gained from using the approximators is discussed in this paper. The FLOPS code is reviewed. Analysis models are developed and validated for each approximator. The regression method appears to hug the data points, while the neural network approximation follows a mean path. For an analysis cycle, the approximate model required milliseconds of central processing unit (CPU) time versus seconds by the FLOPS code. Performance of the approximators was satisfactory for aircraft analysis. A design optimization capability has been created by coupling the derived analyzers to the optimization test bed CometBoards. The approximators were efficient reanalysis tools in the aircraft design optimization. Instability encountered in the FLOPS analyzer was eliminated. The convergence characteristics were improved for the design optimization. The CPU time required to calculate the optimum solution, measured in hours with the FLOPS code was reduced to minutes with the neural network approximation and to seconds with the regression method. Generation of the approximators required the manipulation of a very large quantity of data. Design sensitivity with respect to the bounds of aircraft constraints is easily generated.
Multidisciplinary design and optimization (MDO) methodology for the aircraft conceptual design
NASA Astrophysics Data System (ADS)
Iqbal, Liaquat Ullah
An integrated design and optimization methodology has been developed for the conceptual design of an aircraft. The methodology brings higher fidelity Computer Aided Design, Engineering and Manufacturing (CAD, CAE and CAM) Tools such as CATIA, FLUENT, ANSYS and SURFCAM into the conceptual design by utilizing Excel as the integrator and controller. The approach is demonstrated to integrate with many of the existing low to medium fidelity codes such as the aerodynamic panel code called CMARC and sizing and constraint analysis codes, thus providing the multi-fidelity capabilities to the aircraft designer. The higher fidelity design information from the CAD and CAE tools for the geometry, aerodynamics, structural and environmental performance is provided for the application of the structured design methods such as the Quality Function Deployment (QFD) and the Pugh's Method. The higher fidelity tools bring the quantitative aspects of a design such as precise measurements of weight, volume, surface areas, center of gravity (CG) location, lift over drag ratio, and structural weight, as well as the qualitative aspects such as external geometry definition, internal layout, and coloring scheme early in the design process. The performance and safety risks involved with the new technologies can be reduced by modeling and assessing their impact more accurately on the performance of the aircraft. The methodology also enables the design and evaluation of the novel concepts such as the blended (BWB) and the hybrid wing body (HWB) concepts. Higher fidelity computational fluid dynamics (CFD) and finite element analysis (FEA) allow verification of the claims for the performance gains in aerodynamics and ascertain risks of structural failure due to different pressure distribution in the fuselage as compared with the tube and wing design. The higher fidelity aerodynamics and structural models can lead to better cost estimates that help reduce the financial risks as well. This helps in
NASA Technical Reports Server (NTRS)
Patniak, Surya N.; Guptill, James D.; Hopkins, Dale A.; Lavelle, Thomas M.
1998-01-01
Nonlinear mathematical-programming-based design optimization can be an elegant method. However, the calculations required to generate the merit function, constraints, and their gradients, which are frequently required, can make the process computational intensive. The computational burden can be greatly reduced by using approximating analyzers derived from an original analyzer utilizing neural networks and linear regression methods. The experience gained from using both of these approximation methods in the design optimization of a high speed civil transport aircraft is the subject of this paper. The Langley Research Center's Flight Optimization System was selected for the aircraft analysis. This software was exercised to generate a set of training data with which a neural network and a regression method were trained, thereby producing the two approximating analyzers. The derived analyzers were coupled to the Lewis Research Center's CometBoards test bed to provide the optimization capability. With the combined software, both approximation methods were examined for use in aircraft design optimization, and both performed satisfactorily. The CPU time for solution of the problem, which had been measured in hours, was reduced to minutes with the neural network approximation and to seconds with the regression method. Instability encountered in the aircraft analysis software at certain design points was also eliminated. On the other hand, there were costs and difficulties associated with training the approximating analyzers. The CPU time required to generate the input-output pairs and to train the approximating analyzers was seven times that required for solution of the problem.
NASA Technical Reports Server (NTRS)
Welstead, Jason; Crouse, Gilbert L., Jr.
2014-01-01
Empirical sizing guidelines such as tail volume coefficients have long been used in the early aircraft design phases for sizing stabilizers, resulting in conservatively stable aircraft. While successful, this results in increased empty weight, reduced performance, and greater procurement and operational cost relative to an aircraft with optimally sized surfaces. Including flight dynamics in the conceptual design process allows the design to move away from empirical methods while implementing modern control techniques. A challenge of flight dynamics and control is the numerous design variables, which are changing fluidly throughout the conceptual design process, required to evaluate the system response to some disturbance. This research focuses on addressing that challenge not by implementing higher order tools, such as computational fluid dynamics, but instead by linking the lower order tools typically used within the conceptual design process so each discipline feeds into the other. In thisresearch, flight dynamics and control was incorporated into the conceptual design process along with the traditional disciplines of vehicle sizing, weight estimation, aerodynamics, and performance. For the controller, a linear quadratic regulator structure with constant gains has been specified to reduce the user input. Coupling all the disciplines in the conceptual design phase allows the aircraft designer to explore larger design spaces where stabilizers are sized according to dynamic response constraints rather than historical static margin and volume coefficient guidelines.
NASA Technical Reports Server (NTRS)
Sobieszczanski-Sobieski, Jaroslaw
1988-01-01
Optimization by decomposition, complex system sensitivity analysis, and a rapid growth of disciplinary sensitivity analysis are some of the recent developments that hold promise of a quantum jump in the support engineers receive from computers in the quantitative aspects of design. Review of the salient points of these techniques is given and illustrated by examples from aircraft design as a process that combines the best of human intellect and computer power to manipulate data.
NASA Technical Reports Server (NTRS)
Sobieszczanski-Sobieski, Jaroslaw
1988-01-01
Optimization by decomposition, complex system sensitivity analysis, and a rapid growth of disciplinary sensitivity analysis are some of the recent developments that hold promise of a quantum jump in the support engineers receive from computers in the quantitative aspects of design. Review of the salient points of these techniques is given and illustrated by examples from aircraft design as a process that combines the best of human intellect and computer power to manipulate data.
Neural Network and Regression Methods Demonstrated in the Design Optimization of a Subsonic Aircraft
NASA Technical Reports Server (NTRS)
Hopkins, Dale A.; Lavelle, Thomas M.; Patnaik, Surya
2003-01-01
The neural network and regression methods of NASA Glenn Research Center s COMETBOARDS design optimization testbed were used to generate approximate analysis and design models for a subsonic aircraft operating at Mach 0.85 cruise speed. The analytical model is defined by nine design variables: wing aspect ratio, engine thrust, wing area, sweep angle, chord-thickness ratio, turbine temperature, pressure ratio, bypass ratio, fan pressure; and eight response parameters: weight, landing velocity, takeoff and landing field lengths, approach thrust, overall efficiency, and compressor pressure and temperature. The variables were adjusted to optimally balance the engines to the airframe. The solution strategy included a sensitivity model and the soft analysis model. Researchers generated the sensitivity model by training the approximators to predict an optimum design. The trained neural network predicted all response variables, within 5-percent error. This was reduced to 1 percent by the regression method. The soft analysis model was developed to replace aircraft analysis as the reanalyzer in design optimization. Soft models have been generated for a neural network method, a regression method, and a hybrid method obtained by combining the approximators. The performance of the models is graphed for aircraft weight versus thrust as well as for wing area and turbine temperature. The regression method followed the analytical solution with little error. The neural network exhibited 5-percent maximum error over all parameters. Performance of the hybrid method was intermediate in comparison to the individual approximators. Error in the response variable is smaller than that shown in the figure because of a distortion scale factor. The overall performance of the approximators was considered to be satisfactory because aircraft analysis with NASA Langley Research Center s FLOPS (Flight Optimization System) code is a synthesis of diverse disciplines: weight estimation, aerodynamic
NASA Technical Reports Server (NTRS)
Jules, Kenol; Lin, Paul P.
2002-01-01
This paper reviews some of the recent applications of artificial neural networks taken from various works performed by the authors over the last four years at the NASA Glenn Research Center. This paper focuses mainly on two areas. First, artificial neural networks application in design and optimization of aircraft/engine propulsion systems to shorten the overall design cycle. Out of that specific application, a generic design tool was developed, which can be used for most design optimization process. Second, artificial neural networks application in monitoring the microgravity quality onboard the International Space Station, using on-board accelerometers for data acquisition. These two different applications are reviewed in this paper to show the broad applicability of artificial intelligence in various disciplines. The intent of this paper is not to give in-depth details of these two applications, but to show the need to combine different artificial intelligence techniques or algorithms in order to design an optimized or versatile system.
An overview of the Douglas Aircraft Company Aeroelastic Design Optimization Program (ADOP)
NASA Technical Reports Server (NTRS)
Dodd, Alan J.
1989-01-01
From a program manager's viewpoint, the history, scope and architecture of a major structural design program at Douglas Aircraft Company called Aeroelastic Design Optimization Program (ADOP) are described. ADOP was originally intended for the rapid, accurate, cost-effective evaluation of relatively small structural models at the advanced design level, resulting in improved proposal competitiveness and avoiding many costly changes later in the design cycle. Before release of the initial version in November 1987, however, the program was expanded to handle very large production-type analyses.
The multidisciplinary design optimization of a distributed propulsion blended-wing-body aircraft
NASA Astrophysics Data System (ADS)
Ko, Yan-Yee Andy
The purpose of this study is to examine the multidisciplinary design optimization (MDO) of a distributed propulsion blended-wing-body (BWB) aircraft. The BWB is a hybrid shape resembling a flying wing, placing the payload in the inboard sections of the wing. The distributed propulsion concept involves replacing a small number of large engines with many smaller engines. The distributed propulsion concept considered here ducts part of the engine exhaust to exit out along the trailing edge of the wing. The distributed propulsion concept affects almost every aspect of the BWB design. Methods to model these effects and integrate them into an MDO framework were developed. The most important effect modeled is the impact on the propulsive efficiency. There has been conjecture that there will be an increase in propulsive efficiency when there is blowing out of the trailing edge of a wing. A mathematical formulation was derived to explain this. The formulation showed that the jet 'fills in' the wake behind the body, improving the overall aerodynamic/propulsion system, resulting in an increased propulsive efficiency. The distributed propulsion concept also replaces the conventional elevons with a vectored thrust system for longitudinal control. An extension of Spence's Jet Flap theory was developed to estimate the effects of this vectored thrust system on the aircraft longitudinal control. It was found to provide a reasonable estimate of the control capability of the aircraft. An MDO framework was developed, integrating all the distributed propulsion effects modeled. Using a gradient based optimization algorithm, the distributed propulsion BWB aircraft was optimized and compared with a similarly optimized conventional BWB design. Both designs are for an 800 passenger, 0.85 cruise Mach number and 7000 nmi mission. The MDO results found that the distributed propulsion BWB aircraft has a 4% takeoff gross weight and a 2% fuel weight. Both designs have similar planform shapes
NASA Technical Reports Server (NTRS)
Quinlan, Jesse R.; Gern, Frank H.
2016-01-01
Simultaneously achieving the fuel consumption and noise reduction goals set forth by NASA's Environmentally Responsible Aviation (ERA) project requires innovative and unconventional aircraft concepts. In response, advanced hybrid wing body (HWB) aircraft concepts have been proposed and analyzed as a means of meeting these objectives. For the current study, several HWB concepts were analyzed using the Hybrid wing body Conceptual Design and structural optimization (HCDstruct) analysis code. HCDstruct is a medium-fidelity finite element based conceptual design and structural optimization tool developed to fill the critical analysis gap existing between lower order structural sizing approaches and detailed, often finite element based sizing methods for HWB aircraft concepts. Whereas prior versions of the tool used a half-model approach in building the representative finite element model, a full wing-tip-to-wing-tip modeling capability was recently added to HCDstruct, which alleviated the symmetry constraints at the model centerline in place of a free-flying model and allowed for more realistic center body, aft body, and wing loading and trim response. The latest version of HCDstruct was applied to two ERA reference cases, including the Boeing Open Rotor Engine Integration On an HWB (OREIO) concept and the Boeing ERA-0009H1 concept, and results agreed favorably with detailed Boeing design data and related Flight Optimization System (FLOPS) analyses. Following these benchmark cases, HCDstruct was used to size NASA's ERA HWB concepts and to perform a related scaling study.
NASA Technical Reports Server (NTRS)
Welstead, Jason
2014-01-01
This research focused on incorporating stability and control into a multidisciplinary de- sign optimization on a Boeing 737-class advanced concept called the D8.2b. A new method of evaluating the aircraft handling performance using quantitative evaluation of the sys- tem to disturbances, including perturbations, continuous turbulence, and discrete gusts, is presented. A multidisciplinary design optimization was performed using the D8.2b transport air- craft concept. The con guration was optimized for minimum fuel burn using a design range of 3,000 nautical miles. Optimization cases were run using xed tail volume coecients, static trim constraints, and static trim and dynamic response constraints. A Cessna 182T model was used to test the various dynamic analysis components, ensuring the analysis was behaving as expected. Results of the optimizations show that including stability and con- trol in the design process drastically alters the optimal design, indicating that stability and control should be included in conceptual design to avoid system level penalties later in the design process.
A KBE-enabled design framework for cost/weight optimization study of aircraft composite structures
NASA Astrophysics Data System (ADS)
Wang, H.; La Rocca, G.; van Tooren, M. J. L.
2014-10-01
Traditionally, minimum weight is the objective when optimizing airframe structures. This optimization, however, does not consider the manufacturing cost which actually determines the profit of the airframe manufacturer. To this purpose, a design framework has been developed able to perform cost/weight multi-objective optimization of an aircraft component, including large topology variations of the structural configuration. The key element of the proposed framework is a dedicated knowledge based engineering (KBE) application, called multi-model generator, which enables modelling very different product configurations and variants and extract all data required to feed the weight and cost estimation modules, in a fully automated fashion. The weight estimation method developed in this research work uses Finite Element Analysis to calculate the internal stresses of the structural elements and an analytical composite plate sizing method to determine their minimum required thicknesses. The manufacturing cost estimation module was developed on the basis of a cost model available in literature. The capability of the framework was successfully demonstrated by designing and optimizing the composite structure of a business jet rudder. The study case indicates the design framework is able to find the Pareto optimal set for minimum structural weight and manufacturing costin a very quick way. Based on the Pareto set, the rudder manufacturer is in conditions to conduct both internal trade-off studies between minimum weight and minimum cost solutions, as well as to offer the OEM a full set of optimized options to choose, rather than one feasible design.
NASA Technical Reports Server (NTRS)
Hopkins, Dale A.; Patnaik, Surya N.
2000-01-01
A preliminary aircraft engine design methodology is being developed that utilizes a cascade optimization strategy together with neural network and regression approximation methods. The cascade strategy employs different optimization algorithms in a specified sequence. The neural network and regression methods are used to approximate solutions obtained from the NASA Engine Performance Program (NEPP), which implements engine thermodynamic cycle and performance analysis models. The new methodology is proving to be more robust and computationally efficient than the conventional optimization approach of using a single optimization algorithm with direct reanalysis. The methodology has been demonstrated on a preliminary design problem for a novel subsonic turbofan engine concept that incorporates a wave rotor as a cycle-topping device. Computations of maximum thrust were obtained for a specific design point in the engine mission profile. The results (depicted in the figure) show a significant improvement in the maximum thrust obtained using the new methodology in comparison to benchmark solutions obtained using NEPP in a manual design mode.
NASA Astrophysics Data System (ADS)
Yang, Weizhu; Yue, Zhufeng; Li, Lei; Wang, Peiyan
2016-01-01
An optimization procedure combining an automated finite element modelling (AFEM) technique with a ground structure approach (GSA) is proposed for structural layout and sizing design of aircraft wings. The AFEM technique, based on CATIA VBA scripting and PCL programming, is used to generate models automatically considering the arrangement of inner systems. GSA is used for local structural topology optimization. The design procedure is applied to a high-aspect-ratio wing. The arrangement of the integral fuel tank, landing gear and control surfaces is considered. For the landing gear region, a non-conventional initial structural layout is adopted. The positions of components, the number of ribs and local topology in the wing box and landing gear region are optimized to obtain a minimum structural weight. Constraints include tank volume, strength, buckling and aeroelastic parameters. The results show that the combined approach leads to a greater weight saving, i.e. 26.5%, compared with three additional optimizations based on individual design approaches.
NASA Technical Reports Server (NTRS)
Stepner, D. E.; Mehra, R. K.
1973-01-01
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.
Entropy based design and topology optimization for multifunctional composite aircraft skins
NASA Astrophysics Data System (ADS)
Ahlers, Kristopher Henry Otto
The research presented in this dissertation investigates the design methodology of a new integrated radar aircraft skin, which serves a multifunctional role by accommodating structural and thermal loads. Mechanically the skin must be capable of handling tensile and shear loads. The Entropy Generation Minimization (EGM) method is applied to the multifunctional aircraft skin in order to provide a thermodynamic metric for the analysis of different skin material systems. Thermoelectric (TE) devices, sandwiched between the skin and the radar devices for the purpose of harvesting some of the waste heat to be rejected to the ambient, are considered in the analysis. The optimal entropy generation configuration is one that corresponds to the maximum energy harvesting condition, while satisfying tensile and shear running loads. This optimal condition is found to be when the skin is at its thinnest, with a maximum of 65% volume fraction of carbon fiber inclusions. Additionally it is shown that despite increasing the mass of the skin, the overall mission entropy generated decreases. A topology optimization technique is used to determine the optimal material distribution within the multifunctional skin composite at the 65% volume fraction of inclusion material. A three-dimensional design space is created in which thermal and mechanical loads are applied, and optimized using the Solid Isotropic Microstructures with Penalization (SIMP) method. A weighting variable is introduced to allow for scaling of the incident loads, such that either the thermal or mechanical loading condition may be the dominant load affecting the design space. A thermal dominant loading condition produces a topology representative of a domed structure. A mechanically dominate load results in a 'truss' shaped structure. Novel topologies between these two extreme solutions are realized and presented. Materials research was conducted on carbon foams to assess their viability as a multifunctional skin material
Recent experience with multidisciplinary analysis and optimization in advanced aircraft design
NASA Technical Reports Server (NTRS)
Dollyhigh, Samuel M.; Sobieszczanski-Sobieski, Jaroslaw
1990-01-01
The task of modern aircraft design has always been complicated due to the number of intertwined technical factors from the various engineering disciplines. Furthermore, this complexity has been rapidly increasing by the development of such technologies as aeroelasticity tailored materials and structures, active control systems, integrated propulsion/airframe controls, thrust vectoring, and so on. Successful designs that achieve maximum advantage from these new technologies require a thorough understanding of the physical phenomena and the interactions among these phenomena. A study commissioned by the Aeronautical Sciences and Evaluation Board of the National Research Council has gone so far as to identify technology integration as a new discipline from which many future aeronautical advancements will arise. Regardless of whether one considers integration as a new discipline or not, it is clear to all engineers involved in aircraft design and analysis that better methods are required. In the past, designers conducted parametric studies in which a relatively small number of principal characteristics were varied to determine the effect on design requirements which were themselves often diverse and contradictory. Once a design was chosen, it then passed through the various engineers' disciplines whose principal task was to make the chosen design workable. Working in a limited design space, the discipline expert sometimes improved the concept, but more often than not, the result was in the form of a penalty to make the original concept workable. If an insurmountable problem was encountered, the process began over. Most design systems that attempt to account for disciplinary interactions have large empirical elements and reliance on past experience is a poor guide in obtaining maximum utilizations of new technologies. Further compounding the difficulty of design is that as the aeronautical sciences have matured, the discipline specialist's area of research has generally
NASA Astrophysics Data System (ADS)
Arrieta, Albert Joseph
2001-07-01
Damage tolerance analysis (DTA) was considered in the global design optimization of an aircraft wing structure. Residual strength and fatigue life requirements, based on the damage tolerance philosophy, were investigated as new design constraints. In general, accurate fatigue prediction is difficult if the load environment is not known with a high degree of certainty. To address this issue, a probabilistic approach was used to describe the uncertain load environment. Probabilistic load spectra models were developed from flight recorder data. The global/local finite element approach allowed local fatigue requirements to be considered in the global design optimization. AFGROW fatigue crack growth analysis provided a new strength criterion for satisfying damage tolerance requirements within a global optimization environment. Initial research with the ASTROS program used the probabilistic load model and this damage tolerance constraint to optimize cracked skin panels on the lower wing of a fighter/attack aircraft. For an aerodynamic and structural model similar to an F-16, ASTROS simulated symmetric and asymmetric maneuvers during the optimization. Symmetric maneuvers, without underwing stores, produced the highest stresses and drove the optimization of the inboard lower wing skin. Asymmetric maneuvers, with underwing stores, affected the optimum thickness of the outboard hard points. Subsequent design optimizations included von Mises stress, aileron effectiveness, and lift effectiveness constraints simultaneously. This optimization was driven by the DTA and von Mises stress constraints and, therefore, DTA requirements can have an active role to play in preliminary aircraft design.
A design procedure for the handling qualities optimization of the X-29A aircraft
NASA Technical Reports Server (NTRS)
Bosworth, John T.; Cox, Timothy H.
1989-01-01
The techniques used to improve the pitch-axis handling qualities of the X-29A wing-canard-planform fighter aircraft are reviewed. The aircraft and its FCS are briefly described, and the design method, which works within the existing FCS architecture, is characterized in detail. Consideration is given to the selection of design goals and design variables, the definition and calculation of the cost function, the validation of the mathematical model on the basis of flight-test data, and the validation of the improved design by means of nonlinear simulations. Flight tests of the improved design are shown to verify the simulation results.
NASA Technical Reports Server (NTRS)
Alkamhawi, Hani; Greiner, Tom; Fuerst, Gerry; Luich, Shawn; Stonebraker, Bob; Wray, Todd
1990-01-01
A hypersonic aircraft is designed which uses scramjets to accelerate from Mach 6 to Mach 10 and sustain that speed for two minutes. Different propulsion systems were considered and it was decided that the aircraft would use one full scale turbofan-ramjet. Two solid rocket boosters were added to save fuel and help the aircraft pass through the transonic region. After considering aerodynamics, aircraft design, stability and control, cooling systems, mission profile, and landing systems, a conventional aircraft configuration was chosen over that of a waverider. The conventional design was chosen due to its landing characteristics and the relative expense compared to the waverider. Fuel requirements and the integration of the engine systems and their inlets are also taken into consideration in the final design. A hypersonic aircraft was designed which uses scramjets to accelerate from Mach 6 to Mach 10 and sustain that speed for two minutes. Different propulsion systems were considered and a full scale turbofan-ramjet was chosen. Two solid rocket boosters were added to save fuel and help the aircraft pass through the transonic reqion. After the aerodynamics, aircraft design, stability and control, cooling systems, mission profile, landing systems, and their physical interactions were considered, a conventional aircraft configuration was chosen over that of a waverider. The conventional design was chosen due to its landing characteristics and the relative expense compared to the waverider. Fuel requirements and the integration of the engine systems and their inlets were also considered in the designing process.
A design procedure for the handling qualities optimization of the X-29A aircraft
NASA Technical Reports Server (NTRS)
Bosworth, John T.; Cox, Timothy H.
1989-01-01
A design technique for handling qualities improvement was developed for the X-29A aircraft. As with any new aircraft, the X-29A control law designers were presented with a relatively high degree of uncertainty in their mathematical models. The presence of uncertainties, and the high level of static instability of the X-29A caused the control law designers to stress stability and robustness over handling qualities. During flight test, the mathematical models of the vehicle were validated or corrected to match the vehicle dynamic behavior. The updated models were then used to fine tune the control system to provide fighter-like handling characteristics. A design methodology was developed which works within the existing control system architecture to provide improved handling qualities and acceptable stability with a minimum of cost in both implementation as well as software verification and validation.
Stochastic Methods for Aircraft Design
NASA Technical Reports Server (NTRS)
Pelz, Richard B.; Ogot, Madara
1998-01-01
The global stochastic optimization method, simulated annealing (SA), was adapted and applied to various problems in aircraft design. The research was aimed at overcoming the problem of finding an optimal design in a space with multiple minima and roughness ubiquitous to numerically generated nonlinear objective functions. SA was modified to reduce the number of objective function evaluations for an optimal design, historically the main criticism of stochastic methods. SA was applied to many CFD/MDO problems including: low sonic-boom bodies, minimum drag on supersonic fore-bodies, minimum drag on supersonic aeroelastic fore-bodies, minimum drag on HSCT aeroelastic wings, FLOPS preliminary design code, another preliminary aircraft design study with vortex lattice aerodynamics, HSR complete aircraft aerodynamics. In every case, SA provided a simple, robust and reliable optimization method which found optimal designs in order 100 objective function evaluations. Perhaps most importantly, from this academic/industrial project, technology has been successfully transferred; this method is the method of choice for optimization problems at Northrop Grumman.
Adaptive Flight Control Design with Optimal Control Modification on an F-18 Aircraft Model
NASA Technical Reports Server (NTRS)
Burken, John J.; Nguyen, Nhan T.; Griffin, Brian J.
2010-01-01
In the presence of large uncertainties, a control system needs to be able to adapt rapidly to regain performance. Fast adaptation is referred to as the implementation of adaptive control with a large adaptive gain to reduce the tracking error rapidly; however, a large adaptive gain can lead to high-frequency oscillations which can adversely affect the robustness of an adaptive control law. A new adaptive control modification is presented that can achieve robust adaptation with a large adaptive gain without incurring high-frequency oscillations as with the standard model-reference adaptive control. The modification is based on the minimization of the Y2 norm of the tracking error, which is formulated as an optimal control problem. The optimality condition is used to derive the modification using the gradient method. The optimal control modification results in a stable adaptation and allows a large adaptive gain to be used for better tracking while providing sufficient robustness. A damping term (v) is added in the modification to increase damping as needed. Simulations were conducted on a damaged F-18 aircraft (McDonnell Douglas, now The Boeing Company, Chicago, Illinois) with both the standard baseline dynamic inversion controller and the adaptive optimal control modification technique. The results demonstrate the effectiveness of the proposed modification in tracking a reference model.
NASA Technical Reports Server (NTRS)
Lan, C. Edward; Ge, Fuying
1989-01-01
Control system design for general nonlinear flight dynamic models is considered through numerical simulation. The design is accomplished through a numerical optimizer coupled with analysis of flight dynamic equations. The general flight dynamic equations are numerically integrated and dynamic characteristics are then identified from the dynamic response. The design variables are determined iteratively by the optimizer to optimize a prescribed objective function which is related to desired dynamic characteristics. Generality of the method allows nonlinear effects to aerodynamics and dynamic coupling to be considered in the design process. To demonstrate the method, nonlinear simulation models for an F-5A and an F-16 configurations are used to design dampers to satisfy specifications on flying qualities and control systems to prevent departure. The results indicate that the present method is simple in formulation and effective in satisfying the design objectives.
NASA Astrophysics Data System (ADS)
Gagnon, Hugo
This thesis represents a step forward to bring geometry parameterization and control on par with the disciplinary analyses involved in shape optimization, particularly high-fidelity aerodynamic shape optimization. Central to the proposed methodology is the non-uniform rational B-spline, used here to develop a new geometry generator and geometry control system applicable to the aerodynamic design of both conventional and unconventional aircraft. The geometry generator adopts a component-based approach, where any number of predefined but modifiable (parametric) wing, fuselage, junction, etc., components can be arbitrarily assembled to generate the outer mold line of aircraft geometry. A unique Python-based user interface incorporating an interactive OpenGL windowing system is proposed. Together, these tools allow for the generation of high-quality, C2 continuous (or higher), and customized aircraft geometry with fast turnaround. The geometry control system tightly integrates shape parameterization with volume mesh movement using a two-level free-form deformation approach. The framework is augmented with axial curves, which are shown to be flexible and efficient at parameterizing wing systems of arbitrary topology. A key aspect of this methodology is that very large shape deformations can be achieved with only a few, intuitive control parameters. Shape deformation consumes a few tenths of a second on a single processor and surface sensitivities are machine accurate. The geometry control system is implemented within an existing aerodynamic optimizer comprising a flow solver for the Euler equations and a sequential quadratic programming optimizer. Gradients are evaluated exactly with discrete-adjoint variables. The algorithm is first validated by recovering an elliptical lift distribution on a rectangular wing, and then demonstrated through the exploratory shape optimization of a three-pronged feathered winglet leading to a span efficiency of 1.22 under a height
Advanced hypersonic aircraft design
NASA Technical Reports Server (NTRS)
Utzinger, Rob; Blank, Hans-Joachim; Cox, Craig; Harvey, Greg; Mckee, Mike; Molnar, Dave; Nagy, Greg; Petersen, Steve
1992-01-01
The objective of this design project is to develop the hypersonic reconnaissance aircraft to replace the SR-71 and to complement existing intelligence gathering devices. The initial design considerations were to create a manned vehicle which could complete its mission with at least two airborne refuelings. The aircraft must travel between Mach 4 and Mach 7 at an altitude of 80,000 feet for a maximum range of 12,000 nautical miles. The vehicle should have an air breathing propulsion system at cruise. With a crew of two, the aircraft should be able to take off and land on a 10,000 foot runway, and the yearly operational costs were not to exceed $300 million. Finally, the aircraft should exhibit stealth characteristics, including a minimized radar cross-section (RCS) and a reduced sonic boom. The technology used in this vehicle should allow for production between the years 1993 and 1995.
Optimal design of control systems for prevention of aircraft flight departure
NASA Technical Reports Server (NTRS)
Lan, C. Edward; Ge, Fu-Ying
1990-01-01
Control system design for general nonlinear flight dynamic models is considered through numerical simulation. The design is accomplished through a numerical optimizer coupled with analysis of flight dynamic equations. In the analysis, the general nonlinear flight dynamic equations with nonlinear aerodynamics are numerically integrated; and the dynamic characteristics needed in the optimization process are then identified from the dynamic response. To assure a reasonable solution, the initial input values of the design variables are estimated through a sensitivity analysis. To demonstrate the method, the pitch departure prevention for an F-16 configuration is demonstrated.
NASA Technical Reports Server (NTRS)
Azzano, Christopher P.
1992-01-01
Control of a large jet transport aircraft without the use of conventional control surfaces was studied. Engine commands were used to attempt to recreate the forces and moments typically provided by the elevator, ailerons, and rudder. Necessary conditions for aircraft controllability were developed pertaining to aircraft configuration such as the number of engines and engine placement. An optimal linear quadratic regulator controller was developed for the Boeing 707-720, in particular, for regulation of its natural dynamic modes. The design used a method of assigning relative weights to the natural modes, i.e., phugoid and dutch roll, for a more intuitive selection of the cost function. A prototype pilot command interface was then integrated into the loop based on pseudorate command of both pitch and roll. Closed loop dynamics were evaluated first with a batch linear simulation and then with a real time high fidelity piloted simulation. The NASA research pilots assisted in evaluation of closed loop handling qualities for typical cruise and landing tasks. Recommendations for improvement on this preliminary study of optimal propulsion only flight control are provided.
Aircraft digital control design methods
NASA Technical Reports Server (NTRS)
Powell, J. D.; Parsons, E.; Tashker, M. G.
1976-01-01
Variations in design methods for aircraft digital flight control are evaluated and compared. The methods fall into two categories; those where the design is done in the continuous domain (or s plane) and those where the design is done in the discrete domain (or z plane). Design method fidelity is evaluated by examining closed loop root movement and the frequency response of the discretely controlled continuous aircraft. It was found that all methods provided acceptable performance for sample rates greater than 10 cps except the uncompensated s plane design method which was acceptable above 20 cps. A design procedure based on optimal control methods was proposed that provided the best fidelity at very slow sample rates and required no design iterations for changing sample rates.
NASA Technical Reports Server (NTRS)
Lores, M. E.; Smith, P. R.; Large, R. A.
1980-01-01
An efficient transonic wing design procedure based upon numerical optimization together with three-dimensional transonic methods has been developed and used to design an advanced transport wing. The method development included an examination of the use of both full potential and extended small disturbance analysis codes and demonstrated that the former formulation was more reliable. In either case, the design procedure is economical and easy to use. Design verification in a unique semi-span test arrangement demonstrated that the design method produced a wing which satisfied the study design requirements. However, aeroelastic deformation of the wing occurred during the wind tunnel test. The computational methods used in the design procedure were employed to assess the effect of the aeroelastic deformation. The paper concludes with an evaluation of the design procedure and recommendation for its improvement.
NASA Technical Reports Server (NTRS)
Garg, Sanjay; Schmidt, Phillip H.
1993-01-01
A parameter optimization framework has earlier been developed to solve the problem of partitioning a centralized controller into a decentralized, hierarchical structure suitable for integrated flight/propulsion control implementation. This paper presents results from the application of the controller partitioning optimization procedure to IFPC design for a Short Take-Off and Vertical Landing (STOVL) aircraft in transition flight. The controller partitioning problem and the parameter optimization algorithm are briefly described. Insight is provided into choosing various 'user' selected parameters in the optimization cost function such that the resulting optimized subcontrollers will meet the characteristics of the centralized controller that are crucial to achieving the desired closed-loop performance and robustness, while maintaining the desired subcontroller structure constraints that are crucial for IFPC implementation. The optimization procedure is shown to improve upon the initial partitioned subcontrollers and lead to performance comparable to that achieved with the centralized controller. This application also provides insight into the issues that should be addressed at the centralized control design level in order to obtain implementable partitioned subcontrollers.
Aircraft Optimization for Minimum Environmental Impact
NASA Technical Reports Server (NTRS)
Antoine, Nicolas; Kroo, Ilan M.
2001-01-01
The objective of this research is to investigate the tradeoff between operating cost and environmental acceptability of commercial aircraft. This involves optimizing the aircraft design and mission to minimize operating cost while constraining exterior noise and emissions. Growth in air traffic and airport neighboring communities has resulted in increased pressure to severely penalize airlines that do not meet strict local noise and emissions requirements. As a result, environmental concerns have become potent driving forces in commercial aviation. Traditionally, aircraft have been first designed to meet performance and cost goals, and adjusted to satisfy the environmental requirements at given airports. The focus of the present study is to determine the feasibility of including noise and emissions constraints in the early design of the aircraft and mission. This paper introduces the design tool and results from a case study involving a 250-passenger airliner.
NASA Astrophysics Data System (ADS)
Shrivastava, Sachin; Mohite, P. M.
2015-01-01
The minimization of weight and maximization of payload is an ever challenging design procedure for air vehicles. The present study has been carried out with an objective to redesign control surface of an advanced all-metallic fighter aircraft. In this study, the structure made up of high strength aluminum, titanium and ferrous alloys has been attempted to replace by carbon fiber composite (CFC) skin, ribs and stiffeners. This study presents an approach towards development of a methodology for optimization of first-ply failure index (FI) in unidirectional fibrous laminates using Genetic-Algorithms (GA) under quasi-static loading. The GAs, by the application of its operators like reproduction, cross-over, mutation and elitist strategy, optimize the ply-orientations in laminates so as to have minimum FI of Tsai-Wu first-ply failure criterion. The GA optimization procedure has been implemented in MATLAB and interfaced with commercial software ABAQUS using python scripting. FI calculations have been carried out in ABAQUS with user material subroutine (UMAT). The GA's application gave reasonably well-optimized ply-orientations combination at a faster convergence rate. However, the final optimized sequence of ply-orientations is obtained by tweaking the sequences given by GA's based on industrial practices and experience, whenever needed. The present study of conversion of an all metallic structure to partial CFC structure has led to 12% of weight reduction. Therefore, the approach proposed here motivates designer to use CFC with a confidence.
NASA Technical Reports Server (NTRS)
1997-01-01
Successful commercialization of the AirCraft SYNThesis (ACSYNT) tool has resulted in the creation of Phoenix Integration, Inc. ACSYNT has been exclusively licensed to the company, an outcome of a seven year, $3 million effort to provide unique software technology to a focused design engineering market. Ames Research Center formulated ACSYNT and in working with the Virginia Polytechnic Institute CAD Laboratory, began to design and code a computer-aided design for ACSYNT. Using a Joint Sponsored Research Agreement, Ames formed an industry-government-university alliance to improve and foster research and development for the software. As a result of the ACSYNT Institute, the software is becoming a predominant tool for aircraft conceptual design. ACSYNT has been successfully applied to high- speed civil transport configuration, subsonic transports, and supersonic fighters.
Optimization in fractional aircraft ownership
NASA Astrophysics Data System (ADS)
Septiani, R. D.; Pasaribu, H. M.; Soewono, E.; Fayalita, R. A.
2012-05-01
Fractional Aircraft Ownership is a new concept in flight ownership management system where each individual or corporation may own a fraction of an aircraft. In this system, the owners have privilege to schedule their flight according to their needs. Fractional management companies (FMC) manages all aspects of aircraft operations, including utilization of FMC's aircraft in combination of outsourced aircrafts. This gives the owners the right to enjoy the benefits of private aviations. However, FMC may have complicated business requirements that neither commercial airlines nor charter airlines faces. Here, optimization models are constructed to minimize the number of aircrafts in order to maximize the profit and to minimize the daily operating cost. In this paper, three kinds of demand scenarios are made to represent different flight operations from different types of fractional owners. The problems are formulated as an optimization of profit and a daily operational cost to find the optimum flight assignments satisfying the weekly and daily demand respectively from the owners. Numerical results are obtained by Genetic Algorithm method.
NASA Technical Reports Server (NTRS)
Riedel, S. A.
1979-01-01
A method by which modern and classical control theory techniques may be integrated in a synergistic fashion and used in the design of practical flight control systems is presented. A general procedure is developed, and several illustrative examples are included. Emphasis is placed not only on the synthesis of the design, but on the assessment of the results as well. The first step is to establish the differences, distinguishing characteristics and connections between the modern and classical control theory approaches. Ultimately, this uncovers a relationship between bandwidth goals familiar in classical control and cost function weights in the equivalent optimal system. In order to obtain a practical optimal solution, it is also necessary to formulate the problem very carefully, and each choice of state, measurement and output variable must be judiciously considered. Once design goals are established and problem formulation completed, the control system is synthesized in a straightforward manner. Three steps are involved: filter-observer solution, regulator solution, and the combination of those two into the controller. Assessment of the controller permits and examination and expansion of the synthesis results.
The design of sport and touring aircraft
NASA Technical Reports Server (NTRS)
Eppler, R.; Guenther, W.
1984-01-01
General considerations concerning the design of a new aircraft are discussed, taking into account the objective to develop an aircraft can satisfy economically a certain spectrum of tasks. Requirements related to the design of sport and touring aircraft included in the past mainly a high cruising speed and short take-off and landing runs. Additional requirements for new aircraft are now low fuel consumption and optimal efficiency. A computer program for the computation of flight performance makes it possible to vary automatically a number of parameters, such as flight altitude, wing area, and wing span. The appropriate design characteristics are to a large extent determined by the selection of the flight altitude. Three different wing profiles are compared. Potential improvements with respect to the performance of the aircraft and its efficiency are related to the use of fiber composites, the employment of better propeller profiles, more efficient engines, and the utilization of suitable instrumentation for optimal flight conduction.
Conceptual design optimization study
NASA Technical Reports Server (NTRS)
Hollowell, S. J.; Beeman, E. R., II; Hiyama, R. M.
1990-01-01
The feasibility of applying multilevel functional decomposition and optimization techniques to conceptual design of advanced fighter aircraft was investigated. Applying the functional decomposition techniques to the conceptual design phase appears to be feasible. The initial implementation of the modified design process will optimize wing design variables. A hybrid approach, combining functional decomposition techniques for generation of aerodynamic and mass properties linear sensitivity derivatives with existing techniques for sizing mission performance and optimization, is proposed.
Design Considerations for Laminar Flow Control Aircraft
NASA Technical Reports Server (NTRS)
Sturgeon, R. F.; Bennett, J. A.
1976-01-01
A study was conducted to investigate major design considerations involved in the application of laminar flow control to the wings and empennage of long range subsonic transport aircraft compatible with initial operation in 1985. For commercial transports with a design mission range of 10,186 km (5500 n mil) and a payload of 200 passengers, parametric configuration analyses were conducted to evaluate the effect of aircraft performance, operational, and geometric parameters on fuel efficiency. Study results indicate that major design goals for aircraft optimization include maximization of aspect ratio and wing loading and minimization of wing sweep consistent with wing volume and airport performance requirements.
Design of Ultra-High-Power-Density Machine Optimized for Future Aircraft
NASA Technical Reports Server (NTRS)
Choi, Benjamin B.
2004-01-01
The NASA Glenn Research Center's Structural Mechanics and Dynamics Branch is developing a compact, nonpolluting, bearingless electric machine with electric power supplied by fuel cells for future "more-electric" aircraft with specific power in the projected range of 50 hp/lb, whereas conventional electric machines generate usually 0.2 hp/lb. The use of such electric drives for propulsive fans or propellers depends on the successful development of ultra-high-power-density machines. One possible candidate for such ultra-high-power-density machines, a round-rotor synchronous machine with an engineering current density as high as 20,000 A/sq cm, was selected to investigate how much torque and power can be produced.
A system approach to aircraft optimization
NASA Technical Reports Server (NTRS)
Sobieszczanski-Sobieski, Jaroslaw
1991-01-01
Mutual couplings among the mathematical models of physical phenomena and parts of a system such as an aircraft complicate the design process because each contemplated design change may have a far reaching consequence throughout the system. Techniques are outlined for computing these influences as system design derivatives useful for both judgemental and formal optimization purposes. The techniques facilitate decomposition of the design process into smaller, more manageable tasks and they form a methodology that can easily fit into existing engineering organizations and incorporate their design tools.
Practical optimal flight control system design for helicopter aircraft. Volume 1: Technical Report
NASA Technical Reports Server (NTRS)
Hofmann, L. G.; Riedel, S. A.; Mcruer, D.
1980-01-01
A method by which modern and classical theory techniques may be integrated in a synergistic fashion and used in the design of practical flight control systems is presented. A general procedure is developed, and several illustrative examples are included. Emphasis is placed not only on the synthesis of the design, but on the assessment of the results as well.
Control optimization, stabilization and computer algorithms for aircraft applications
NASA Technical Reports Server (NTRS)
1975-01-01
Research related to reliable aircraft design is summarized. Topics discussed include systems reliability optimization, failure detection algorithms, analysis of nonlinear filters, design of compensators incorporating time delays, digital compensator design, estimation for systems with echoes, low-order compensator design, descent-phase controller for 4-D navigation, infinite dimensional mathematical programming problems and optimal control problems with constraints, robust compensator design, numerical methods for the Lyapunov equations, and perturbation methods in linear filtering and control.
NASA Technical Reports Server (NTRS)
Maccready, P. B.; Lissaman, P. B. S.; Morgan, W. R.; Burke, J. D.
1981-01-01
Two piloted aircraft have been developed and flown powered solely by photovoltaic cells in a program sponsored by the DuPont Company. The 30.8-kg (68-lb), 21.6-m (71-ft) span, Gossamer Penguin was used as a solar test bed, making a 2.6-km (1.6-mile) flight in August 1980. The 88.1-kg (194-lb), 14.3-m (47-ft) span Solar Challenger was developed for long flights in normal turbulence. Stressed to +9 G, it utilizes Kevlar, Nomex honeycomb-graphite sandwich wall tubes, expanded polystyrene foam ribs, and Mylar skin. With a 54.9-kg (121-lb) airframe, 33.1-kg (73-lb) propulsion system, and a 45.4-kg (100-lb) pilot, it flies on 1400 watts. In summer, the projected maximum climb is 1.0 m/s (200 ft/min) at 9,150 m (30,000 ft). Sixty purely solar-powered flights were made during winter 1980-1981. Using thermals, 1,070 m (3,500 ft) was reached with 115-minute duration.
Innovations in Aircraft Design
NASA Technical Reports Server (NTRS)
1997-01-01
The Boeing 777 carries with it basic and applied research, technology, and aerodynamic knowledge honed at several NASA field centers. Several Langley Research Center innovations instrumental to the development of the aircraft include knowledge of how to reduce engine and other noise for passengers and terminal residents, increased use of lightweight aerospace composite structures for increased fuel efficiency and range, and wind tunnel tests confirming the structural integrity of 777 wing-airframe integration. Test results from Marshall Space Flight Center aimed at improving the performance of the Space Shuttle engines led to improvements in the airplane's new, more efficient jet engines. Finally, fostered by Ames Research Center, the Boeing 777 blankets that protect areas of the plane from high temperatures and fire have a lineage to Advanced Flexible Reusable Surface Insulation used on certain areas of the Space Shuttle. According to Boeing Company estimates, the 777 has captured three-quarters of new orders for airplanes in its class since the program was launched.
Progress in aircraft design since 1903
NASA Technical Reports Server (NTRS)
1974-01-01
Significant developments in aviation history are documented to show the advancements in aircraft design which have taken place since 1903. Each aircraft is identified according to the manufacturer, powerplant, dimensions, normal weight, and typical performance. A narrative summary of the major accomplishments of the aircraft is provided. Photographs of each aircraft are included.
Aircraft empennage structural detail design
NASA Technical Reports Server (NTRS)
Meholic, Greg; Brown, Rhonda; Hall, Melissa; Harvey, Robert; Singer, Michael; Tella, Gustavo
1993-01-01
This project involved the detailed design of the aft fuselage and empennage structure, vertical stabilizer, rudder, horizontal stabilizer, and elevator for the Triton primary flight trainer. The main design goals under consideration were to illustrate the integration of the control systems devices used in the tail surfaces and their necessary structural supports as well as the elevator trim, navigational lighting system, electrical systems, tail-located ground tie, and fuselage/cabin interface structure. Accommodations for maintenance, lubrication, adjustment, and repairability were devised. Weight, fabrication, and (sub)assembly goals were addressed. All designs were in accordance with the FAR Part 23 stipulations for a normal category aircraft.
Control optimization, stabilization and computer algorithms for aircraft applications
NASA Technical Reports Server (NTRS)
Athans, M. (Editor); Willsky, A. S. (Editor)
1982-01-01
The analysis and design of complex multivariable reliable control systems are considered. High performance and fault tolerant aircraft systems are the objectives. A preliminary feasibility study of the design of a lateral control system for a VTOL aircraft that is to land on a DD963 class destroyer under high sea state conditions is provided. Progress in the following areas is summarized: (1) VTOL control system design studies; (2) robust multivariable control system synthesis; (3) adaptive control systems; (4) failure detection algorithms; and (5) fault tolerant optimal control theory.
Aircraft digital control design methods
NASA Technical Reports Server (NTRS)
Tashker, M. G.; Powell, J. D.
1975-01-01
Investigations were conducted in two main areas: the first area is control system design, and the goals were to define the limits of 'digitized S-Plane design techniques' vs. sample rate, to show the results of a 'direct digital design technique', and to compare the two methods; the second area was to evaluate the roughness of autopilot designs parametrically versus sample rate. Goals of the first area were addressed by (1) an analysis of a 2nd order example using both design methods, (2) a linear analysis of the complete 737 aircraft with an autoland obtained using the digitized S-plane technique, (3) linear analysis of a high frequency 737 approximation with the autoland from a direct digital design technique, and (4) development of a simulation for evaluation of the autopilots with disturbances and nonlinearities included. Roughness evaluation was studied by defining an experiment to be carried out on the Langley motion simulator and coordinated with analysis at Stanford.
TRACON Aircraft Arrival Planning and Optimization Through Spatial Constraint Satisfaction
NASA Technical Reports Server (NTRS)
Bergh, Christopher P.; Krzeczowski, Kenneth J.; Davis, Thomas J.; Denery, Dallas G. (Technical Monitor)
1995-01-01
A new aircraft arrival planning and optimization algorithm has been incorporated into the Final Approach Spacing Tool (FAST) in the Center-TRACON Automation System (CTAS) developed at NASA-Ames Research Center. FAST simulations have been conducted over three years involving full-proficiency, level five air traffic controllers from around the United States. From these simulations an algorithm, called Spatial Constraint Satisfaction, has been designed, coded, undergone testing, and soon will begin field evaluation at the Dallas-Fort Worth and Denver International airport facilities. The purpose of this new design is an attempt to show that the generation of efficient and conflict free aircraft arrival plans at the runway does not guarantee an operationally acceptable arrival plan upstream from the runway -information encompassing the entire arrival airspace must be used in order to create an acceptable aircraft arrival plan. This new design includes functions available previously but additionally includes necessary representations of controller preferences and workload, operationally required amounts of extra separation, and integrates aircraft conflict resolution. As a result, the Spatial Constraint Satisfaction algorithm produces an optimized aircraft arrival plan that is more acceptable in terms of arrival procedures and air traffic controller workload. This paper discusses the current Air Traffic Control arrival planning procedures, previous work in this field, the design of the Spatial Constraint Satisfaction algorithm, and the results of recent evaluations of the algorithm.
Aircraft wing structure detail design
NASA Technical Reports Server (NTRS)
Sager, Garrett L.; Roberts, Ron; Mallon, Bob; Alameri, Mohamed; Steinbach, Bill
1993-01-01
The provisions of this project call for the design of the structure of the wing and carry-through structure for the Viper primary trainer, which is to be certified as a utility category trainer under FAR part 23. The specific items to be designed in this statement of work were Front Spar, Rear Spar, Aileron Structure, Wing Skin, and Fuselage Carry-through Structure. In the design of these parts, provisions for the fuel system, electrical system, and control routing were required. Also, the total weight of the entire wing planform could not exceed 216 lbs. Since this aircraft is to be used as a primary trainer, and the SOW requires a useful life of 107 cycles, it was decided that all of the principle stresses in the structural members would be kept below 10 ksi. The only drawback to this approach is a weight penalty.
Rapid Parameterization Schemes for Aircraft Shape Optimization
NASA Technical Reports Server (NTRS)
Li, Wu
2012-01-01
A rapid shape parameterization tool called PROTEUS is developed for aircraft shape optimization. This tool can be applied directly to any aircraft geometry that has been defined in PLOT3D format, with the restriction that each aircraft component must be defined by only one data block. PROTEUS has eight types of parameterization schemes: planform, wing surface, twist, body surface, body scaling, body camber line, shifting/scaling, and linear morphing. These parametric schemes can be applied to two types of components: wing-type surfaces (e.g., wing, canard, horizontal tail, vertical tail, and pylon) and body-type surfaces (e.g., fuselage, pod, and nacelle). These schemes permit the easy setup of commonly used shape modification methods, and each customized parametric scheme can be applied to the same type of component for any configuration. This paper explains the mathematics for these parametric schemes and uses two supersonic configurations to demonstrate the application of these schemes.
The Cal Poly aircraft design program
NASA Technical Reports Server (NTRS)
Sandlin, Doral R.; Van't Riet, Robert
1993-01-01
Discussed is the aircraft design program at Cal Poly, SLO. The history of the program and the impact of the NASA/USRA ADP are presented. Examples of student design accomplishments are included. Questions on how the aircraft design education process can be improved are postulated.
Information fusion based optimal control for large civil aircraft system.
Zhen, Ziyang; Jiang, Ju; Wang, Xinhua; Gao, Chen
2015-03-01
Wind disturbance has a great influence on landing security of Large Civil Aircraft. Through simulation research and engineering experience, it can be found that PID control is not good enough to solve the problem of restraining the wind disturbance. This paper focuses on anti-wind attitude control for Large Civil Aircraft in landing phase. In order to improve the riding comfort and the flight security, an information fusion based optimal control strategy is presented to restrain the wind in landing phase for maintaining attitudes and airspeed. Data of Boeing707 is used to establish a nonlinear mode with total variables of Large Civil Aircraft, and then two linear models are obtained which are divided into longitudinal and lateral equations. Based on engineering experience, the longitudinal channel adopts PID control and C inner control to keep longitudinal attitude constant, and applies autothrottle system for keeping airspeed constant, while an information fusion based optimal regulator in the lateral control channel is designed to achieve lateral attitude holding. According to information fusion estimation, by fusing hard constraint information of system dynamic equations and the soft constraint information of performance index function, optimal estimation of the control sequence is derived. Based on this, an information fusion state regulator is deduced for discrete time linear system with disturbance. The simulation results of nonlinear model of aircraft indicate that the information fusion optimal control is better than traditional PID control, LQR control and LQR control with integral action, in anti-wind disturbance performance in the landing phase. PMID:25440950
Hodograph analysis in aircraft trajectory optimization
NASA Technical Reports Server (NTRS)
Cliff, Eugene M.; Seywald, Hans; Bless, Robert R.
1993-01-01
An account is given of key geometrical concepts involved in the use of a hodograph as an optimal control theory resource which furnishes a framework for geometrical interpretation of the minimum principle. Attention is given to the effects of different convexity properties on the hodograph, which bear on the existence of solutions and such types of controls as chattering controls, 'bang-bang' control, and/or singular control. Illustrative aircraft trajectory optimization problems are examined in view of this use of the hodograph.
Aircraft optimization by a system approach: Achievements and trends
NASA Technical Reports Server (NTRS)
Sobieszczanski-Sobieski, Jaroslaw
1992-01-01
Recently emerging methodology for optimal design of aircraft treated as a system of interacting physical phenomena and parts is examined. The methodology is found to coalesce into methods for hierarchic, non-hierarchic, and hybrid systems all dependent on sensitivity analysis. A separate category of methods has also evolved independent of sensitivity analysis, hence suitable for discrete problems. References and numerical applications are cited. Massively parallel computer processing is seen as enabling technology for practical implementation of the methodology.
Designing for aircraft structural crashworthiness
NASA Technical Reports Server (NTRS)
Thomson, R. G.; Caiafa, C.
1981-01-01
This report describes structural aviation crash dynamics research activities being conducted on general aviation aircraft and transport aircraft. The report includes experimental and analytical correlations of load-limiting subfloor and seat configurations tested dynamically in vertical drop tests and in a horizontal sled deceleration facility. Computer predictions using a finite-element nonlinear computer program, DYCAST, of the acceleration time-histories of these innovative seat and subfloor structures are presented. Proposed application of these computer techniques, and the nonlinear lumped mass computer program KRASH, to transport aircraft crash dynamics is discussed. A proposed FAA full-scale crash test of a fully instrumented radio controlled transport airplane is also described.
Sonic Boom Mitigation Through Aircraft Design and Adjoint Methodology
NASA Technical Reports Server (NTRS)
Rallabhandi, Siriam K.; Diskin, Boris; Nielsen, Eric J.
2012-01-01
This paper presents a novel approach to design of the supersonic aircraft outer mold line (OML) by optimizing the A-weighted loudness of sonic boom signature predicted on the ground. The optimization process uses the sensitivity information obtained by coupling the discrete adjoint formulations for the augmented Burgers Equation and Computational Fluid Dynamics (CFD) equations. This coupled formulation links the loudness of the ground boom signature to the aircraft geometry thus allowing efficient shape optimization for the purpose of minimizing the impact of loudness. The accuracy of the adjoint-based sensitivities is verified against sensitivities obtained using an independent complex-variable approach. The adjoint based optimization methodology is applied to a configuration previously optimized using alternative state of the art optimization methods and produces additional loudness reduction. The results of the optimizations are reported and discussed.
Alloy design for aircraft engines
NASA Astrophysics Data System (ADS)
Pollock, Tresa M.
2016-08-01
Metallic materials are fundamental to advanced aircraft engines. While perceived as mature, emerging computational, experimental and processing innovations are expanding the scope for discovery and implementation of new metallic materials for future generations of advanced propulsion systems.
NASA Technical Reports Server (NTRS)
Sobel, K.; Kaufman, H.
1979-01-01
A stochastic linear model that accounts for process parameter and initial uncertainty, measurement noise, and a restricted number of measurable outputs was used to determine feedback gains useful for reducing the vertical acceleration which results from the presence of a vertical wind gust. Considered in the study were the influence of various feedback configurations, the effects of sensor noise, flight condition changes, and initialization procedures. Results showed that for sixth order linearized longitudinal motion, a controller with feedback on three states could be designed for effective gust alleviation taking into account both sensor noise and flight condition variation.
Optimal control of a low wing-loading STOL aircraft
NASA Technical Reports Server (NTRS)
Cunningham, T. B.
1976-01-01
Linear optimal quadratic control theory is applied to a low wing-loading STOL aircraft for ride quality and flight path following. Design criteria include minimum rms response to wind turbulence and desired transient response characteristics. Design techniques include proper choosing of design versus evaluation models, choosing appropriate performance index responses, and use of classical evaluation techniques. Results are obtained through a combination of frequency response shaping and gust observation. Effects of control rate and authority saturation are examined with a new rapid calculation of random input describing functions. Parameter sensitivity is also evaluated using a Liapunov type matrix equation.
Structural design of supersonic cruise aircraft
NASA Technical Reports Server (NTRS)
Fischler, J. E.
1976-01-01
The major efforts leading to an efficient structural design include: (1) the analysis methods used to improve the structural model optimization and compare the structural concepts, (2) the analysis and description of the fail-safe, crack growth, and residual strength studies and tests, (3) baseline structural trade studies to determine optimum structural weights including effects of geometry changes, strength, fail-safety, aeroelastics and flutter, 6AL-4V annealed titanium in structural efficiency after 70,000 hours at temperature, (5) the study of three structural models for aircraft at 2.0 Mach, 2.2 Mach, and 2.4 Mach cruise speeds, (6) the study of many structural concepts to determine their weight efficiencies; and (7) the determination of the requirements for large-scale structural development testing.
Flight director design for a STOL aircraft.
NASA Technical Reports Server (NTRS)
Seitz, W. R.; Goodson, R. E.
1973-01-01
The study described here was intended to explore some ways of giving the pilot the information necessary to perform as well as an autopilot. A fixed-base flight simulator was built to study pilot/director/aircraft performance. Instrument rated pilots used the different director-displays in an approach down to touchdown, including a flare and decrab maneuver. Two well established areas of control theory are combined in the design of the cockpit display. Optimal control theory and the theory of manual control are used to find the feedback gains required to drive the display symbols. Conclusions based on the simulator are presented. The results show that the director display developed in the work provides more than adequate information for simulated landing in highly turbulent conditions.
The design of four hypersonic reconnaissance aircraft
NASA Technical Reports Server (NTRS)
Gregorek, G. M.; Detwiler, D. T.
1992-01-01
Four different hypersonic reconnaissance aircraft were designed by separate student teams. These aircraft were designed to provide the U.S. with a system to acquire aerial tactical reconnaissance when satellite reconnaissance proved unobtainable or ineffective. The design requirements given for this project stated that these aircraft must carry a 7500 lb, 250 cu ft payload of electronic and photographic intelligence gathering equipment over a target area at speeds between Mach 4-7 and at altitudes above 80,000 ft. Two of the aircraft were required to be manned by a crew of two and have a range of 12,000 nmi. One of these was to use airborne refueling to complete its mission while the other was not to use any refueling. The other two aircraft were required to be unmanned with a range of 6,000 nmi. One of these was to take off from another aircraft. The final details of all four aircraft designs along with an overview of the design process is provided.
Supersonic aircraft optimization for minimizing drag and sonic boom
NASA Astrophysics Data System (ADS)
Chan, Martin Kia-Yin
A design tool incorporating classical sonic boom theory, computational fluid dynamics and a multi-objective genetic algorithm was developed for low-boom supersonic aircraft optimization. Both sonic boom and drag were optimized simultaneously and a Pareto optimal set of designs ranging from minimum boom to minimum drag was obtained for each optimization. Since sonic boom was optimized directly, the method had broader applicability than the traditional inverse method. A high-order three-dimensional panel method was used for sonic boom prediction. The traditional linear source model was fast but did not account for wing-body aerodynamic interaction. Euler solutions were expensive for computing sonic booms because a large number of grid points were needed to accurately predict the pressure signature away from the aircraft. For the Mach number and configurations of interest, the panel code showed good agreement with Euler but at a fraction of the cost. A loudness metric was shown to have advantages over initial overpressure and peak overpressure for measuring shaped sonic booms. However, optimization results obtained using calculated loudness raised concerns about the robustness of the solution to atmospheric disturbance. Peak overpressure minimization also produced reasonable sonic boom signatures and appeared more robust to atmospheric uncertainties, but the resulting loudness was not as good. Better convergence was also observed with peak overpressure. Two supersonic business jets were optimized. One was a conventional configuration; the other was a natural laminar flow configuration. Optimization results obtained using loudness and peak overpressure were compared. A non-axisymmetric fuselage was optimized and found to reduce the inviscid drag by 9 to 30 percent at the same sonic boom loudness.
A technique for integrating engine cycle and aircraft configuration optimization
NASA Technical Reports Server (NTRS)
Geiselhart, Karl A.
1994-01-01
A method for conceptual aircraft design that incorporates the optimization of major engine design variables for a variety of cycle types was developed. The methodology should improve the lengthy screening process currently involved in selecting an appropriate engine cycle for a given application or mission. The new capability will allow environmental concerns such as airport noise and emissions to be addressed early in the design process. The ability to rapidly perform optimization and parametric variations using both engine cycle and aircraft design variables, and to see the impact on the aircraft, should provide insight and guidance for more detailed studies. A brief description of the aircraft performance and mission analysis program and the engine cycle analysis program that were used is given. A new method of predicting propulsion system weight and dimensions using thermodynamic cycle data, preliminary design, and semi-empirical techniques is introduced. Propulsion system performance and weights data generated by the program are compared with industry data and data generated using well established codes. The ability of the optimization techniques to locate an optimum is demonstrated and some of the problems that had to be solved to accomplish this are illustrated. Results from the application of the program to the analysis of three supersonic transport concepts installed with mixed flow turbofans are presented. The results from the application to a Mach 2.4, 5000 n.mi. transport indicate that the optimum bypass ratio is near 0.45 with less than 1 percent variation in minimum gross weight for bypass ratios ranging from 0.3 to 0.6. In the final application of the program, a low sonic boom fix a takeoff gross weight concept that would fly at Mach 2.0 overwater and at Mach 1.6 overland is compared with a baseline concept of the same takeoff gross weight that would fly Mach 2.4 overwater and subsonically overland. The results indicate that for the design mission
Low Reynolds number, long endurance aircraft design
Foch, R.J.; Ailinger, K.G. )
1992-02-01
Airplanes are typically designed to maximize range at the highest practical cruising speed. However, several missions require extended duration rather than range, and favor the slowest possible cruise speed. Such missions include surveillance, radio relay, and ship's electronic decoy. These missions are ideally suited for advanced technology unmanned aircraft, either remotely piloted or autonomous. Feasibility studies have been conducted and flight demonstrator prototypes of such unique aircraft have been under steady research and development at the Naval Research Laboratory since 1978. This paper discusses the design aspects and tradeoffs unique to small, slow speed long endurance unmanned aircraft operating at wing chord Reynolds numbers between 150,000 and 500,000. Additionally, many of these low Reynolds number-driven design features have applicability to high altitude, long endurance aircraft. 6 refs.
NASA Astrophysics Data System (ADS)
Haley, Stephen
Aircraft manufacturer Aquavion Systems is currently designing and constructing prototypes for its revolutionary new fleet of aircraft called the SeaStryder. During the prototyping phase, it was discovered that the center-of-gravity of the SeaStryder600 was too far aft and outside of the acceptable range. To solve this design issue, it was hypothesized that the weight of the horizontal stabilizer may be reduced without compromising its structural integrity. The following document analyzes this hypothesis and provides two alternative designs. Each design exceeds the design requirements, meets additional requirements requested by industry, and provides a significant degree of weight savings. The first design provides a 25% weight reduction. The second design provides an 18% weight reduction as well as a 160% increase in loading capacity. The designs proposed have both been verified through the use of Finite Element Analysis as well as by means of experimentation where two prototype wings were constructed and tested to failure confirming the analytical results.
Modification of ACSYNT aircraft computer program for preliminary design
NASA Technical Reports Server (NTRS)
Biezad, Daniel J.; Rojos-Oviedo, Ruben
1994-01-01
This paper presents the development of a computer simulation of agility flight test techniques. Its purpose is to evaluate the agility of aircraft configurations early in the preliminary design phase. The simulation module is integrated into the NASA Ames aircraft synthesis design code. Trade studies using the agility module embedded within the design code to simulate the combat cycle time agility metric are illustrated using a Northrop F-20 aircraft model. Results show that the agility module is effective in analyzing the influence of common parameters such as thrust-to-weight ratio and wing loading on agility criteria. The module can also compare the agility potential between different configurations and has the capability to optimize agility performance early in the design process.
Cascade Optimization Strategy for Aircraft and Air-Breathing Propulsion System Concepts
NASA Technical Reports Server (NTRS)
Patnaik, Surya N.; Lavelle, Thomas M.; Hopkins, Dale A.; Coroneos, Rula M.
1996-01-01
Design optimization for subsonic and supersonic aircraft and for air-breathing propulsion engine concepts has been accomplished by soft-coupling the Flight Optimization System (FLOPS) and the NASA Engine Performance Program analyzer (NEPP), to the NASA Lewis multidisciplinary optimization tool COMETBOARDS. Aircraft and engine design problems, with their associated constraints and design variables, were cast as nonlinear optimization problems with aircraft weight and engine thrust as the respective merit functions. Because of the diversity of constraint types and the overall distortion of the design space, the most reliable single optimization algorithm available in COMETBOARDS could not produce a satisfactory feasible optimum solution. Some of COMETBOARDS' unique features, which include a cascade strategy, variable and constraint formulations, and scaling devised especially for difficult multidisciplinary applications, successfully optimized the performance of both aircraft and engines. The cascade method has two principal steps: In the first, the solution initiates from a user-specified design and optimizer, in the second, the optimum design obtained in the first step with some random perturbation is used to begin the next specified optimizer. The second step is repeated for a specified sequence of optimizers or until a successful solution of the problem is achieved. A successful solution should satisfy the specified convergence criteria and have several active constraints but no violated constraints. The cascade strategy available in the combined COMETBOARDS, FLOPS, and NEPP design tool converges to the same global optimum solution even when it starts from different design points. This reliable and robust design tool eliminates manual intervention in the design of aircraft and of air-breathing propulsion engines where it eases the cycle analysis procedures. The combined code is also much easier to use, which is an added benefit. This paper describes COMETBOARDS
Multifidelity Analysis and Optimization for Supersonic Design
NASA Technical Reports Server (NTRS)
Kroo, Ilan; Willcox, Karen; March, Andrew; Haas, Alex; Rajnarayan, Dev; Kays, Cory
2010-01-01
Supersonic aircraft design is a computationally expensive optimization problem and multifidelity approaches over a significant opportunity to reduce design time and computational cost. This report presents tools developed to improve supersonic aircraft design capabilities including: aerodynamic tools for supersonic aircraft configurations; a systematic way to manage model uncertainty; and multifidelity model management concepts that incorporate uncertainty. The aerodynamic analysis tools developed are appropriate for use in a multifidelity optimization framework, and include four analysis routines to estimate the lift and drag of a supersonic airfoil, a multifidelity supersonic drag code that estimates the drag of aircraft configurations with three different methods: an area rule method, a panel method, and an Euler solver. In addition, five multifidelity optimization methods are developed, which include local and global methods as well as gradient-based and gradient-free techniques.
Cascaded Optimization for a Persistent Data Ferrying Unmanned Aircraft
NASA Astrophysics Data System (ADS)
Carfang, Anthony
This dissertation develops and assesses a cascaded method for designing optimal periodic trajectories and link schedules for an unmanned aircraft to ferry data between stationary ground nodes. This results in a fast solution method without the need to artificially constrain system dynamics. Focusing on a fundamental ferrying problem that involves one source and one destination, but includes complex vehicle and Radio-Frequency (RF) dynamics, a cascaded structure to the system dynamics is uncovered. This structure is exploited by reformulating the nonlinear optimization problem into one that reduces the independent control to the vehicle's motion, while the link scheduling control is folded into the objective function and implemented as an optimal policy that depends on candidate motion control. This formulation is proven to maintain optimality while reducing computation time in comparison to traditional ferry optimization methods. The discrete link scheduling problem takes the form of a combinatorial optimization problem that is known to be NP-Hard. A derived necessary condition for optimality guides the development of several heuristic algorithms, specifically the Most-Data-First Algorithm and the Knapsack Adaptation. These heuristics are extended to larger ferrying scenarios, and assessed analytically and through Monte Carlo simulation, showing better throughput performance in the same order of magnitude of computation time in comparison to other common link scheduling policies. The cascaded optimization method is implemented with a novel embedded software system on a small, unmanned aircraft to validate the simulation results with field experiments. To address the sensitivity of results on trajectory tracking performance, a system that combines motion and link control with waypoint-based navigation is developed and assessed through field experiments. The data ferrying algorithms are further extended by incorporating a Gaussian process to opportunistically learn
Multi-level systems modeling and optimization for novel aircraft
NASA Astrophysics Data System (ADS)
Subramanian, Shreyas Vathul
This research combines the disciplines of system-of-systems (SoS) modeling, platform-based design, optimization and evolving design spaces to achieve a novel capability for designing solutions to key aeronautical mission challenges. A central innovation in this approach is the confluence of multi-level modeling (from sub-systems to the aircraft system to aeronautical system-of-systems) in a way that coordinates the appropriate problem formulations at each level and enables parametric search in design libraries for solutions that satisfy level-specific objectives. The work here addresses the topic of SoS optimization and discusses problem formulation, solution strategy, the need for new algorithms that address special features of this problem type, and also demonstrates these concepts using two example application problems - a surveillance UAV swarm problem, and the design of noise optimal aircraft and approach procedures. This topic is critical since most new capabilities in aeronautics will be provided not just by a single air vehicle, but by aeronautical Systems of Systems (SoS). At the same time, many new aircraft concepts are pressing the boundaries of cyber-physical complexity through the myriad of dynamic and adaptive sub-systems that are rising up the TRL (Technology Readiness Level) scale. This compositional approach is envisioned to be active at three levels: validated sub-systems are integrated to form conceptual aircraft, which are further connected with others to perform a challenging mission capability at the SoS level. While these multiple levels represent layers of physical abstraction, each discipline is associated with tools of varying fidelity forming strata of 'analysis abstraction'. Further, the design (composition) will be guided by a suitable hierarchical complexity metric formulated for the management of complexity in both the problem (as part of the generative procedure and selection of fidelity level) and the product (i.e., is the mission
Application of decomposition techniques to the preliminary design of a transport aircraft
NASA Technical Reports Server (NTRS)
Rogan, J. E.; Kolb, M. A.
1987-01-01
A nonlinear constrained optimization problem describing the preliminary design process for a transport aircraft has been formulated. A multifaceted decomposition of the optimization problem has been made. Flight dynamics, flexible aircraft loads and deformations, and preliminary structural design subproblems appear prominently in the decomposition. The use of design process decomposition for scheduling design projects, a new system integration approach to configuration control, and the application of object-centered programming to a new generation of design tools are discussed.
Subsonic Aircraft With Regression and Neural-Network Approximators Designed
NASA Technical Reports Server (NTRS)
Patnaik, Surya N.; Hopkins, Dale A.
2004-01-01
At the NASA Glenn Research Center, NASA Langley Research Center's Flight Optimization System (FLOPS) and the design optimization testbed COMETBOARDS with regression and neural-network-analysis approximators have been coupled to obtain a preliminary aircraft design methodology. For a subsonic aircraft, the optimal design, that is the airframe-engine combination, is obtained by the simulation. The aircraft is powered by two high-bypass-ratio engines with a nominal thrust of about 35,000 lbf. It is to carry 150 passengers at a cruise speed of Mach 0.8 over a range of 3000 n mi and to operate on a 6000-ft runway. The aircraft design utilized a neural network and a regression-approximations-based analysis tool, along with a multioptimizer cascade algorithm that uses sequential linear programming, sequential quadratic programming, the method of feasible directions, and then sequential quadratic programming again. Optimal aircraft weight versus the number of design iterations is shown. The central processing unit (CPU) time to solution is given. It is shown that the regression-method-based analyzer exhibited a smoother convergence pattern than the FLOPS code. The optimum weight obtained by the approximation technique and the FLOPS code differed by 1.3 percent. Prediction by the approximation technique exhibited no error for the aircraft wing area and turbine entry temperature, whereas it was within 2 percent for most other parameters. Cascade strategy was required by FLOPS as well as the approximators. The regression method had a tendency to hug the data points, whereas the neural network exhibited a propensity to follow a mean path. The performance of the neural network and regression methods was considered adequate. It was at about the same level for small, standard, and large models with redundancy ratios (defined as the number of input-output pairs to the number of unknown coefficients) of 14, 28, and 57, respectively. In an SGI octane workstation (Silicon Graphics
NASA Technical Reports Server (NTRS)
Allan, Brian; Owens, Lewis
2010-01-01
In support of the Blended-Wing-Body aircraft concept, a new flow control hybrid vane/jet design has been developed for use in a boundary-layer-ingesting (BLI) offset inlet in transonic flows. This inlet flow control is designed to minimize the engine fan-face distortion levels and the first five Fourier harmonic half amplitudes while maximizing the inlet pressure recovery. This concept represents a potentially enabling technology for quieter and more environmentally friendly transport aircraft. An optimum vane design was found by minimizing the engine fan-face distortion, DC60, and the first five Fourier harmonic half amplitudes, while maximizing the total pressure recovery. The optimal vane design was then used in a BLI inlet wind tunnel experiment at NASA Langley's 0.3-meter transonic cryogenic tunnel. The experimental results demonstrated an 80-percent decrease in DPCPavg, the reduction in the circumferential distortion levels, at an inlet mass flow rate corresponding to the middle of the operational range at the cruise condition. Even though the vanes were designed at a single inlet mass flow rate, they performed very well over the entire inlet mass flow range tested in the wind tunnel experiment with the addition of a small amount of jet flow control. While the circumferential distortion was decreased, the radial distortion on the outer rings at the aerodynamic interface plane (AIP) increased. This was a result of the large boundary layer being distributed from the bottom of the AIP in the baseline case to the outer edges of the AIP when using the vortex generator (VG) vane flow control. Experimental results, as already mentioned, showed an 80-percent reduction of DPCPavg, the circumferential distortion level at the engine fan-face. The hybrid approach leverages strengths of vane and jet flow control devices, increasing inlet performance over a broader operational range with significant reduction in mass flow requirements. Minimal distortion level requirements
14 CFR 183.27 - Designated aircraft maintenance inspectors.
Code of Federal Regulations, 2012 CFR
2012-01-01
... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Designated aircraft maintenance inspectors...: Privileges Â§ 183.27 Designated aircraft maintenance inspectors. A designated aircraft maintenance inspector (DAMI) may approve maintenance on civil aircraft used by United States military flying clubs in...
14 CFR 183.27 - Designated aircraft maintenance inspectors.
Code of Federal Regulations, 2014 CFR
2014-01-01
... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Designated aircraft maintenance inspectors...: Privileges Â§ 183.27 Designated aircraft maintenance inspectors. A designated aircraft maintenance inspector (DAMI) may approve maintenance on civil aircraft used by United States military flying clubs in...
14 CFR 183.27 - Designated aircraft maintenance inspectors.
Code of Federal Regulations, 2013 CFR
2013-01-01
... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Designated aircraft maintenance inspectors...: Privileges Â§ 183.27 Designated aircraft maintenance inspectors. A designated aircraft maintenance inspector (DAMI) may approve maintenance on civil aircraft used by United States military flying clubs in...
14 CFR 183.27 - Designated aircraft maintenance inspectors.
Code of Federal Regulations, 2010 CFR
2010-01-01
... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Designated aircraft maintenance inspectors...: Privileges Â§ 183.27 Designated aircraft maintenance inspectors. A designated aircraft maintenance inspector (DAMI) may approve maintenance on civil aircraft used by United States military flying clubs in...
An integrated computer system for preliminary design of advanced aircraft.
NASA Technical Reports Server (NTRS)
Fulton, R. E.; Sobieszczanski, J.; Landrum, E. J.
1972-01-01
A progress report is given on the first phase of a research project to develop a system of Integrated Programs for Aerospace-Vehicle Design (IPAD) which is intended to automate to the largest extent possible the preliminary and detailed design of advanced aircraft. The approach used is to build a pilot system and simultaneously to carry out two major contractual studies to define a practical IPAD system preparatory to programing. The paper summarizes the specifications and goals of the IPAD system, the progress to date, and any conclusion reached regarding its feasibility and scope. Sample calculations obtained with the pilot system are given for aircraft preliminary designs optimized with respect to discipline parameters, such as weight or L/D, and these results are compared with designs optimized with respect to overall performance parameters, such as range or payload.
Neural Network Prediction of New Aircraft Design Coefficients
NASA Technical Reports Server (NTRS)
Norgaard, Magnus; Jorgensen, Charles C.; Ross, James C.
1997-01-01
This paper discusses a neural network tool for more effective aircraft design evaluations during wind tunnel tests. Using a hybrid neural network optimization method, we have produced fast and reliable predictions of aerodynamical coefficients, found optimal flap settings, and flap schedules. For validation, the tool was tested on a 55% scale model of the USAF/NASA Subsonic High Alpha Research Concept aircraft (SHARC). Four different networks were trained to predict coefficients of lift, drag, moment of inertia, and lift drag ratio (C(sub L), C(sub D), C(sub M), and L/D) from angle of attack and flap settings. The latter network was then used to determine an overall optimal flap setting and for finding optimal flap schedules.
Simulation Packages Expand Aircraft Design Options
NASA Technical Reports Server (NTRS)
2013-01-01
In 2001, NASA released a new approach to computational fluid dynamics that allows users to perform automated analysis on complex vehicle designs. In 2010, Palo Alto, California-based Desktop Aeronautics acquired a license from Ames Research Center to sell the technology. Today, the product assists organizations in the design of subsonic aircraft, space planes, spacecraft, and high speed commercial jets.
Design of a spanloader cargo aircraft
NASA Technical Reports Server (NTRS)
1989-01-01
With a growing demand for fast international freight service, the slow-moving cargo ships currently in use will soon find a substantial portion of their clients looking elsewhere. One candidate for filling this expected gap in the freight market is a span-loading aircraft (or 'flying wing') capable of long-range operation with extremely large payloads. This report summarizes the design features of an aircraft capable of fulfilling a long-haul, high-capacity cargo mission. The spanloader seeks to gain advantage over conventional aircraft by eliminating the aircraft fuselage and thus reducing empty weight. The primary disadvantage of this configuration is that the cargo-containing wing tends to be thick, thus posing a challenge to the airfoil designer. It also suffers from stability and control problems not encountered by conventional aircraft. The result is an interesting, challenging exercise in unconventional design. The report that follows is a student written synopsis of an effort judged to be the best of eight designs developed during the year 1988-1989.
Analysis and design technology for high-speed aircraft structures
NASA Technical Reports Server (NTRS)
Starnes, James H., Jr.; Camarda, Charles J.
1992-01-01
Recent high-speed aircraft structures research activities at NASA Langley Research Center are described. The following topics are covered: the development of analytical and numerical solutions to global and local thermal and structural problems, experimental verification of analysis methods, identification of failure mechanisms, and the incorporation of analysis methods into design and optimization strategies. The paper describes recent NASA Langley advances in analysis and design methods, structural and thermal concepts, and test methods.
Aircraft integrated design and analysis: A classroom experience
NASA Technical Reports Server (NTRS)
1988-01-01
AAE 451 is the capstone course required of all senior undergraduates in the School of Aeronautics and Astronautics at Purdue University. During the past year the first steps of a long evolutionary process were taken to change the content and expectations of this course. These changes are the result of the availability of advanced computational capabilities and sophisticated electronic media availability at Purdue. This presentation will describe both the long range objectives and this year's experience using the High Speed Commercial Transport (HSCT) design, the AIAA Long Duration Aircraft design and a Remotely Piloted Vehicle (RPV) design proposal as project objectives. The central goal of these efforts was to provide a user-friendly, computer-software-based, environment to supplement traditional design course methodology. The Purdue University Computer Center (PUCC), the Engineering Computer Network (ECN), and stand-alone PC's were used for this development. This year's accomplishments centered primarily on aerodynamics software obtained from the NASA Langley Research Center and its integration into the classroom. Word processor capability for oral and written work and computer graphics were also blended into the course. A total of 10 HSCT designs were generated, ranging from twin-fuselage and forward-swept wing aircraft, to the more traditional delta and double-delta wing aircraft. Four Long Duration Aircraft designs were submitted, together with one RPV design tailored for photographic surveillance. Supporting these activities were three video satellite lectures beamed from NASA/Langley to Purdue. These lectures covered diverse areas such as an overview of HSCT design, supersonic-aircraft stability and control, and optimization of aircraft performance. Plans for next year's effort will be reviewed, including dedicated computer workstation utilization, remote satellite lectures, and university/industrial cooperative efforts.
Programs To Optimize Spacecraft And Aircraft Trajectories
NASA Technical Reports Server (NTRS)
Brauer, G. L.; Petersen, F. M.; Cornick, D.E.; Stevenson, R.; Olson, D. W.
1994-01-01
POST/6D POST is set of two computer programs providing ability to target and optimize trajectories of powered or unpowered spacecraft or aircraft operating at or near rotating planet. POST treats point-mass, three-degree-of-freedom case. 6D POST treats more-general rigid-body, six-degree-of-freedom (with point masses) case. Used to solve variety of performance, guidance, and flight-control problems for atmospheric and orbital vehicles. Applications include computation of performance or capability of vehicle in ascent, or orbit, and during entry into atmosphere, simulation and analysis of guidance and flight-control systems, dispersion-type analyses and analyses of loads, general-purpose six-degree-of-freedom simulation of controlled and uncontrolled vehicles, and validation of performance in six degrees of freedom. Written in FORTRAN 77 and C language. Two machine versions available: one for SUN-series computers running SunOS(TM) (LAR-14871) and one for Silicon Graphics IRIS computers running IRIX(TM) operating system (LAR-14869).
Modeling Programs Increase Aircraft Design Safety
NASA Technical Reports Server (NTRS)
2012-01-01
Flutter may sound like a benign word when associated with a flag in a breeze, a butterfly, or seaweed in an ocean current. When used in the context of aerodynamics, however, it describes a highly dangerous, potentially deadly condition. Consider the case of the Lockheed L-188 Electra Turboprop, an airliner that first took to the skies in 1957. Two years later, an Electra plummeted to the ground en route from Houston to Dallas. Within another year, a second Electra crashed. In both cases, all crew and passengers died. Lockheed engineers were at a loss as to why the planes wings were tearing off in midair. For an answer, the company turned to NASA s Transonic Dynamics Tunnel (TDT) at Langley Research Center. At the time, the newly renovated wind tunnel offered engineers the capability of testing aeroelastic qualities in aircraft flying at transonic speeds near or just below the speed of sound. (Aeroelasticity is the interaction between aerodynamic forces and the structural dynamics of an aircraft or other structure.) Through round-the-clock testing in the TDT, NASA and industry researchers discovered the cause: flutter. Flutter occurs when aerodynamic forces acting on a wing cause it to vibrate. As the aircraft moves faster, certain conditions can cause that vibration to multiply and feed off itself, building to greater amplitudes until the flutter causes severe damage or even the destruction of the aircraft. Flutter can impact other structures as well. Famous film footage of the Tacoma Narrows Bridge in Washington in 1940 shows the main span of the bridge collapsing after strong winds generated powerful flutter forces. In the Electra s case, faulty engine mounts allowed a type of flutter known as whirl flutter, generated by the spinning propellers, to transfer to the wings, causing them to vibrate violently enough to tear off. Thanks to the NASA testing, Lockheed was able to correct the Electra s design flaws that led to the flutter conditions and return the
The design of digital-adaptive controllers for VTOL aircraft
NASA Technical Reports Server (NTRS)
Stengel, R. F.; Broussard, J. R.; Berry, P. W.
1976-01-01
Design procedures for VTOL automatic control systems have been developed and are presented. Using linear-optimal estimation and control techniques as a starting point, digital-adaptive control laws have been designed for the VALT Research Aircraft, a tandem-rotor helicopter which is equipped for fully automatic flight in terminal area operations. These control laws are designed to interface with velocity-command and attitude-command guidance logic, which could be used in short-haul VTOL operations. Developments reported here include new algorithms for designing non-zero-set-point digital regulators, design procedures for rate-limited systems, and algorithms for dynamic control trim setting.
Structural analysis at aircraft conceptual design stage
NASA Astrophysics Data System (ADS)
Mansouri, Reza
In the past 50 years, computers have helped by augmenting human efforts with tremendous pace. The aircraft industry is not an exception. Aircraft industry is more than ever dependent on computing because of a high level of complexity and the increasing need for excellence to survive a highly competitive marketplace. Designers choose computers to perform almost every analysis task. But while doing so, existing effective, accurate and easy to use classical analytical methods are often forgotten, which can be very useful especially in the early phases of the aircraft design where concept generation and evaluation demands physical visibility of design parameters to make decisions [39, 2004]. Structural analysis methods have been used by human beings since the very early civilization. Centuries before computers were invented; the pyramids were designed and constructed by Egyptians around 2000 B.C, the Parthenon was built by the Greeks, around 240 B.C, Dujiangyan was built by the Chinese. Persepolis, Hagia Sophia, Taj Mahal, Eiffel tower are only few more examples of historical buildings, bridges and monuments that were constructed before we had any advancement made in computer aided engineering. Aircraft industry is no exception either. In the first half of the 20th century, engineers used classical method and designed civil transport aircraft such as Ford Tri Motor (1926), Lockheed Vega (1927), Lockheed 9 Orion (1931), Douglas DC-3 (1935), Douglas DC-4/C-54 Skymaster (1938), Boeing 307 (1938) and Boeing 314 Clipper (1939) and managed to become airborne without difficulty. Evidencing, while advanced numerical methods such as the finite element analysis is one of the most effective structural analysis methods; classical structural analysis methods can also be as useful especially during the early phase of a fixed wing aircraft design where major decisions are made and concept generation and evaluation demands physical visibility of design parameters to make decisions
Computational aerodynamics applications to transport aircraft design
NASA Technical Reports Server (NTRS)
Henne, P. A.
1983-01-01
Examples are cited in assessing the effect that computational aerodynamics has had on the design of transport aircraft. The application of computational potential flow methods to wing design and to high-lift system design is discussed. The benefits offered by computational aerodynamics in reducing design cost, time, and risk are shown to be substantial.These aerodynamic methods have proved to be particularly effective in exposing inferior or poor aerodynamic designs. Particular attention is given to wing design, where the results have been dramatic.
NASA Technical Reports Server (NTRS)
Cavage, R. L.
1975-01-01
Results are presented of a study of lift-cruise fan V/STOL aircraft for the 1980-1985 time period. Technical and operating characteristics and technology requirements for the ultimate development of this type aircraft are identified. Aircraft individually optimized to perform the antisubmarine warfare, carrier onboard delivery, combat search and rescue, and surveillance and surface attack missions are considered along with a multi-purpose aircraft concept capable of performing all five missions at minimum total program cost. It is shown that lighter and smaller aircraft could be obtained by optimizing the design and fan selection for specific missions.
On a global aerodynamic optimization of a civil transport aircraft
NASA Technical Reports Server (NTRS)
Savu, G.; Trifu, O.
1991-01-01
An aerodynamic optimization procedure developed to minimize the drag to lift ratio of an aircraft configuration: wing - body - tail, in accordance with engineering restrictions, is described. An algorithm developed to search a hypersurface with 18 dimensions, which define an aircraft configuration, is discussed. The results, when considered from the aerodynamic point of view, indicate the optimal configuration is one that combines a lifting fuselage with a canard.
Aerodynamic design trends for commercial aircraft
NASA Technical Reports Server (NTRS)
Hilbig, R.; Koerner, H.
1986-01-01
Recent research on advanced-configuration commercial aircraft at DFVLR is surveyed, with a focus on aerodynamic approaches to improved performance. Topics examined include transonic wings with variable camber or shock/boundary-layer control, wings with reduced friction drag or laminarized flow, prop-fan propulsion, and unusual configurations or wing profiles. Drawings, diagrams, and graphs of predicted performance are provided, and the need for extensive development efforts using powerful computer facilities, high-speed and low-speed wind tunnels, and flight tests of models (mounted on specially designed carrier aircraft) is indicated.
Design of a spanloader cargo aircraft
NASA Technical Reports Server (NTRS)
Weisshaar, Terrence A.
1989-01-01
The design features of an aircraft capable of fulfilling a long haul, high capacity cargo mission are described. This span-loading aircraft, or flying wing, is capable of carrying extremely large payloads and is expected to be in demand to replace the slow-moving cargo ships currently in use. The spanloader seeks to reduce empty weight by eliminating the aircraft fuselage. Disadvantages are the thickness of the cargo-containing wing, and resulting stability and control problems. The spanloader presented here has a small fuselage, low-aspect ratio wings, winglets, and uses six turbofan engines for propulsion. It will have a payload capacity of 300,000 pounds plus 30 first class passengers and 6 crew members. Its projected market is transportation of freight from Europe and the U.S.A. to countries in the Pacific Basin. Cost estimates support its economic feasibility.
Manx: Close air support aircraft preliminary design
NASA Technical Reports Server (NTRS)
Amy, Annie; Crone, David; Hendrickson, Heidi; Willis, Randy; Silva, Vince
1991-01-01
The Manx is a twin engine, twin tailed, single seat close air support design proposal for the 1991 Team Student Design Competition. It blends advanced technologies into a lightweight, high performance design with the following features: High sensitivity (rugged, easily maintained, with night/adverse weather capability); Highly maneuverable (negative static margin, forward swept wing, canard, and advanced avionics result in enhanced aircraft agility); and Highly versatile (design flexibility allows the Manx to contribute to a truly integrated ground team capable of rapid deployment from forward sites).
14 CFR 183.27 - Designated aircraft maintenance inspectors.
Code of Federal Regulations, 2011 CFR
2011-01-01
... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Designated aircraft maintenance inspectors. 183.27 Section 183.27 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF...: Privileges Â§ 183.27 Designated aircraft maintenance inspectors. A designated aircraft maintenance...
Aircraft cabin noise prediction and optimization
NASA Technical Reports Server (NTRS)
Vaicaitis, R.
1985-01-01
Theoretical and experimental studies were conducted to determine the noise transmission into acoustic enclosures ranging from simple rectangular box models to full scale light aircraft in flight. The structural models include simple, stiffened, curved stiffened, and orthotropic panels and double wall windows. The theoretical solutions were obtained by model analysis. Transfer matrix and finite element procedures were utilized. Good agreement between theory and experiment has been achieved. An efficient acoustic add-on treatment was developed for interior noise control in a twin engine light aircraft.
System Synthesis in Preliminary Aircraft Design using Statistical Methods
NASA Technical Reports Server (NTRS)
DeLaurentis, Daniel; Mavris, Dimitri N.; Schrage, Daniel P.
1996-01-01
This paper documents an approach to conceptual and preliminary aircraft design in which system synthesis is achieved using statistical methods, specifically design of experiments (DOE) and response surface methodology (RSM). These methods are employed in order to more efficiently search the design space for optimum configurations. In particular, a methodology incorporating three uses of these techniques is presented. First, response surface equations are formed which represent aerodynamic analyses, in the form of regression polynomials, which are more sophisticated than generally available in early design stages. Next, a regression equation for an overall evaluation criterion is constructed for the purpose of constrained optimization at the system level. This optimization, though achieved in a innovative way, is still traditional in that it is a point design solution. The methodology put forward here remedies this by introducing uncertainty into the problem, resulting a solutions which are probabilistic in nature. DOE/RSM is used for the third time in this setting. The process is demonstrated through a detailed aero-propulsion optimization of a high speed civil transport. Fundamental goals of the methodology, then, are to introduce higher fidelity disciplinary analyses to the conceptual aircraft synthesis and provide a roadmap for transitioning from point solutions to probabalistic designs (and eventually robust ones).
System Synthesis in Preliminary Aircraft Design Using Statistical Methods
NASA Technical Reports Server (NTRS)
DeLaurentis, Daniel; Mavris, Dimitri N.; Schrage, Daniel P.
1996-01-01
This paper documents an approach to conceptual and early preliminary aircraft design in which system synthesis is achieved using statistical methods, specifically Design of Experiments (DOE) and Response Surface Methodology (RSM). These methods are employed in order to more efficiently search the design space for optimum configurations. In particular, a methodology incorporating three uses of these techniques is presented. First, response surface equations are formed which represent aerodynamic analyses, in the form of regression polynomials, which are more sophisticated than generally available in early design stages. Next, a regression equation for an Overall Evaluation Criterion is constructed for the purpose of constrained optimization at the system level. This optimization, though achieved in an innovative way, is still traditional in that it is a point design solution. The methodology put forward here remedies this by introducing uncertainty into the problem, resulting in solutions which are probabilistic in nature. DOE/RSM is used for the third time in this setting. The process is demonstrated through a detailed aero-propulsion optimization of a High Speed Civil Transport. Fundamental goals of the methodology, then, are to introduce higher fidelity disciplinary analyses to the conceptual aircraft synthesis and provide a roadmap for transitioning from point solutions to probabilistic designs (and eventually robust ones).
Preliminary design of a family of close air support aircraft
NASA Technical Reports Server (NTRS)
Tuschhoff, Jeff; Bartel, Rudy; Cox, Brian; Darrah, Paul; Drake, TY; Hendrich, Louis; Hicks, Robin; Holt, Mark; Hoyle, Mark; Kerns, Brian
1989-01-01
A family of three Close Air Support (CAS) aircraft is presented. These aircraft are designed with commonality as the main design objective to reduce the life cycle cost. The aircraft are low wing, twin-boom, pusher turbo-prop configurations. The amount of information displayed to the pilot was reduced to a minimum to greatly simplify the cockpit. The aircraft met the mission specifications and the performance and cost characteristics compared well with other CAS aircraft. The concept of a family of CAS aircraft seems viable after preliminary design.
Applications of structural optimization methods to fixed-wing aircraft and spacecraft in the 1980s
NASA Technical Reports Server (NTRS)
Miura, Hirokazu; Neill, Douglas J.
1992-01-01
This report is the summary of a technical survey on the applications of structural optimization in the U.S. aerospace industry through the 1980s. Since applications to rotary wing aircraft will be covered by other literature, applications to fixed-wing aircraft and spacecraft were considered. It became clear that very significant progress has been made during this decade, indicating this technology is about to become one of the practical tools in computer aided structural design.
Application of trajectory optimization principles to minimize aircraft operating costs
NASA Technical Reports Server (NTRS)
Sorensen, J. A.; Morello, S. A.; Erzberger, H.
1979-01-01
This paper summarizes various applications of trajectory optimization principles that have been or are being devised by both government and industrial researchers to minimize aircraft direct operating costs (DOC). These costs (time and fuel) are computed for aircraft constrained to fly over a fixed range. Optimization theory is briefly outlined, and specific algorithms which have resulted from application of this theory are described. Typical results which demonstrate use of these algorithms and the potential savings which they can produce are given. Finally, need for further trajectory optimization research is presented.
Optimal nonlinear estimation for aircraft flight control in wind shear
NASA Technical Reports Server (NTRS)
Mulgund, Sandeep S.
1994-01-01
The most recent results in an ongoing research effort at Princeton in the area of flight dynamics in wind shear are described. The first undertaking in this project was a trajectory optimization study. The flight path of a medium-haul twin-jet transport aircraft was optimized during microburst encounters on final approach. The assumed goal was to track a reference climb rate during an aborted landing, subject to a minimum airspeed constraint. The results demonstrated that the energy loss through the microburst significantly affected the qualitative nature of the optimal flight path. In microbursts of light to moderate strength, the aircraft was able to track the reference climb rate successfully. In severe microbursts, the minimum airspeed constraint in the optimization forced the aircraft to settle on a climb rate smaller than the target. A tradeoff was forced between the objectives of flight path tracking and stall prevention.
A systematic approach to design for lifelong aircraft evolution
NASA Astrophysics Data System (ADS)
Lim, Dongwook
This research proposes a systematic approach with which the decision makers can evaluate the value and risk of a new aircraft development program, including potential derivative development opportunities. The proposed Evaluation of Lifelong Vehicle Evolution (EvoLVE) method is a two- or multi-stage representation of the aircraft design process that accommodates initial development phases as well as follow-on phases. One of the key elements of this method is the Stochastic Programming with Recourse (SPR) technique, which accounts for uncertainties associated with future requirements. The remedial approach of SPR in its two distinctive problem-solving steps is well suited to aircraft design problems where derivatives, retrofits, and upgrades have been used to fix designs that were once but no longer optimal. The solution approach of SPR is complemented by the Risk-Averse Strategy Selection (RASS) technique to gauge risk associated with vehicle evolution options. In the absence of a full description of the random space, a scenario-based approach captures the randomness with a few probable scenarios and reveals implications of different future events. Last, an interactive framework for decision-making support allows simultaneous navigation of the current and future design space with a greater degree of freedom. A cantilevered beam design problem was set up and solved using the SPR technique to showcase its application to an engineering design setting. The full EvoLVE method was conducted on a notional multi-role fighter based on the F/A-18 Hornet.
Efficient Viscous Design of Realistic Aircraft Configurations
NASA Technical Reports Server (NTRS)
Campbell, Richard L.
2004-01-01
This paper addresses the use of the Constrained Direct Iterative Surface Curvature (CDISC) design method in the aircraft design process. A discussion of some of the requirements for practical use of CFD in the design process is followed by a description of different CFD design methods, along with their relative strengths and weaknesses. A detailed description of the CDISC design method highlights some of the aspects of the method that provide computational efficiency and portability, as well as the flow and geometry constraint capabilities. In addition, an efficient approach to multipoint design, the Weighted Averaging of Geometries (WAG) method, is described and illustrated using a couple of simple examples. The CDISC and WAG methods are then applied to a complex generic business jet geometry using an unstructured grid flow solver to demonstrate the multipoint and multicomponent design capabilities of these methods. Introduction
Survey - Applications of structural optimization methods to fixed wing aircraft and spacecraft
NASA Technical Reports Server (NTRS)
Miura, Hirokazu; Neill, Douglas J.
1992-01-01
Results of a technical survey of the practical applications of structural optimization methods in the U.S. aerospace industry through 1980s are summarized. One of the most important developments in the 80s is the more widespread acceptance of structural optimization as one of the design tools that support practical structural design. Another significant advance is the development of large software tools for production applications. Attention is also given to the tailoring of the computerized design process to the specific environment of each company. The two most important aspects of this tailoring are seamless and easy-to-use incorporation of structural optimization in the overall aerospace design/production process and multidisciplinary integration aimed at ultimate performance optimization of the final product. Some specific applications discussed include the X-29 forward swept wing demonstrator aircraft, composite wing and vertical tail program, fighter wing redesign evaluations, high speed aircraft design, and space structures.
Design Sensitivity for a Subsonic Aircraft Predicted by Neural Network and Regression Models
NASA Technical Reports Server (NTRS)
Hopkins, Dale A.; Patnaik, Surya N.
2005-01-01
A preliminary methodology was obtained for the design optimization of a subsonic aircraft by coupling NASA Langley Research Center s Flight Optimization System (FLOPS) with NASA Glenn Research Center s design optimization testbed (COMETBOARDS with regression and neural network analysis approximators). The aircraft modeled can carry 200 passengers at a cruise speed of Mach 0.85 over a range of 2500 n mi and can operate on standard 6000-ft takeoff and landing runways. The design simulation was extended to evaluate the optimal airframe and engine parameters for the subsonic aircraft to operate on nonstandard runways. Regression and neural network approximators were used to examine aircraft operation on runways ranging in length from 4500 to 7500 ft.
High-order computational fluid dynamics tools for aircraft design.
Wang, Z J
2014-08-13
Most forecasts predict an annual airline traffic growth rate between 4.5 and 5% in the foreseeable future. To sustain that growth, the environmental impact of aircraft cannot be ignored. Future aircraft must have much better fuel economy, dramatically less greenhouse gas emissions and noise, in addition to better performance. Many technical breakthroughs must take place to achieve the aggressive environmental goals set up by governments in North America and Europe. One of these breakthroughs will be physics-based, highly accurate and efficient computational fluid dynamics and aeroacoustics tools capable of predicting complex flows over the entire flight envelope and through an aircraft engine, and computing aircraft noise. Some of these flows are dominated by unsteady vortices of disparate scales, often highly turbulent, and they call for higher-order methods. As these tools will be integral components of a multi-disciplinary optimization environment, they must be efficient to impact design. Ultimately, the accuracy, efficiency, robustness, scalability and geometric flexibility will determine which methods will be adopted in the design process. This article explores these aspects and identifies pacing items. PMID:25024419
High-order computational fluid dynamics tools for aircraft design
Wang, Z. J.
2014-01-01
Most forecasts predict an annual airline traffic growth rate between 4.5 and 5% in the foreseeable future. To sustain that growth, the environmental impact of aircraft cannot be ignored. Future aircraft must have much better fuel economy, dramatically less greenhouse gas emissions and noise, in addition to better performance. Many technical breakthroughs must take place to achieve the aggressive environmental goals set up by governments in North America and Europe. One of these breakthroughs will be physics-based, highly accurate and efficient computational fluid dynamics and aeroacoustics tools capable of predicting complex flows over the entire flight envelope and through an aircraft engine, and computing aircraft noise. Some of these flows are dominated by unsteady vortices of disparate scales, often highly turbulent, and they call for higher-order methods. As these tools will be integral components of a multi-disciplinary optimization environment, they must be efficient to impact design. Ultimately, the accuracy, efficiency, robustness, scalability and geometric flexibility will determine which methods will be adopted in the design process. This article explores these aspects and identifies pacing items. PMID:25024419
Design of a control configured tanker aircraft
NASA Technical Reports Server (NTRS)
Walker, S. A.
1976-01-01
The benefits that accrue from using control configured vehicle (CCV) concepts were examined along with the techniques for applying these concepts to an advanced tanker aircraft design. Reduced static stability (RSS) and flutter mode control (FMC) were the two primary CCV concepts used in the design. The CCV tanker was designed to the same mission requirements specified for a conventional tanker design. A seven degree of freedom mathematical model of the flexible aircraft was derived and used to synthesize a lateral stability augmentation system (SAS), a longitudinal control augmentation system (CAS), and a FMC system. Fatigue life and cost analyses followed the control system synthesis, after which a comparative evaluation of the CCV and conventional tankers was made. This comparison indicated that the CCV weight and cost were lower but that, for this design iteration, the CCV fatigue life was shorter. Also, the CCV crew station acceleration was lower, but the acceleration at the boom operator station was higher relative to the corresponding conventional tanker. Comparison of the design processes used in the CCV and conventional design studies revealed that they were basically the same.
Propulsion Controlled Aircraft design and development
NASA Technical Reports Server (NTRS)
Wells, Edward A.; Urnes, James M., Sr.
1995-01-01
This paper describes the design, development, and ground testing of the propulsion controlled aircraft (PCA) flight control system. A backup flight control system which uses only engine thrust, the PCA system utilizes collective and differential thrust changes to steer an aircraft that experiences partial or complete failure of the hydraulically actuated control surfaces. The objective of the program was to investigate, in flight, the throttles-only control capability of the F-15, using manual control, and also an augmented PCA mode in which computer-controlled thrust was used for flight control. The objective included PCA operation in up-and-away flight and, if performance was adequate, a secondary objective to make actual PCA landings. The PCA design began with a feasibility study which evaluated many control law designs. The study was done using off-line control analysis, simulation, and on-line manned flight simulator tests. Control laws, cockpit displays, and cockpit controls were evaluated by NASA test pilots. A flight test baseline configuration was selected based on projected flight performance, applicability to transport and fighter aircraft, and funding costs. During the PCA software and hardware development, the initial design was updated as data became available from throttle-only flight experiments conducted by NASA on the F-15. This information showed basic airframe characteristics that were not observed in the F-15 flight simulator and resulted in several design changes. After the primary objectives of the PCA flight testing were accomplished, additional PCA modes of operation were developed and implemented. The evolution of the PCA system from the initial feasibility study, control law design, simulation, hardware-in-the-loop tests, pilot-in-the-loop tests, and ground tests is presented.
Optimal and suboptimal control technique for aircraft spin recovery
NASA Technical Reports Server (NTRS)
Young, J. W.
1974-01-01
An analytic investigation has been made of procedures for effecting recovery from equilibrium spin conditions for three assumed aircraft configurations. Three approaches which utilize conventional aerodynamic controls are investigated. Included are a constant control recovery mode, optimal recoveries, and a suboptimal control logic patterned after optimal recovery results. The optimal and suboptimal techniques are shown to yield a significant improvement in recovery performance over that attained by using a constant control recovery procedure.
TRUSS: An intelligent design system for aircraft wings
NASA Technical Reports Server (NTRS)
Bates, Preston R.; Schrage, Daniel P.
1989-01-01
Competitive leadership in the international marketplace, superiority in national defense, excellence in productivity, and safety of both private and public systems are all national defense goals which are dependent on superior engineering design. In recent years, it has become more evident that early design decisions are critical, and when only based on performance often result in products which are too expensive, hard to manufacture, or unsupportable. Better use of computer-aided design tools and information-based technologies is required to produce better quality United States products. A program is outlined here to explore the use of knowledge based expert systems coupled with numerical optimization, database management techniques, and designer interface methods in a networked design environment to improve and assess design changes due to changing emphasis or requirements. The initial structural design of a tiltrotor aircraft wing is used as a representative example to demonstrate the approach being followed.
Aircraft Conceptual Design Using Vehicle Sketch Pad
NASA Technical Reports Server (NTRS)
Fredericks, William J.; Antcliff, Kevin R.; Costa, Guillermo; Deshpande, Nachiket; Moore, Mark D.; Miguel, Edric A. San; Snyder, Alison N.
2010-01-01
Vehicle Sketch Pad (VSP) is a parametric geometry modeling tool that is intended for use in the conceptual design of aircraft. The intent of this software is to rapidly model aircraft configurations without expending the expertise and time that is typically required for modeling with traditional Computer Aided Design (CAD) packages. VSP accomplishes this by using parametrically defined components, such as a wing that is defined by span, area, sweep, taper ratio, thickness to cord, and so on. During this phase of frequent design builds, changes to the model can be rapidly visualized along with the internal volumetric layout. Using this geometry-based approach, parameters such as wetted areas and cord lengths can be easily extracted for rapid external performance analyses, such as a parasite drag buildup. At the completion of the conceptual design phase, VSP can export its geometry to higher fidelity tools. This geometry tool was developed by NASA and is freely available to U.S. companies and universities. It has become integral to conceptual design in the Aeronautics Systems Analysis Branch (ASAB) here at NASA Langley Research Center and is currently being used at over 100 universities, aerospace companies, and other government agencies. This paper focuses on the use of VSP in recent NASA conceptual design studies to facilitate geometry-centered design methodology. Such a process is shown to promote greater levels of creativity, more rapid assessment of critical design issues, and improved ability to quickly interact with higher order analyses. A number of VSP vehicle model examples are compared to CAD-based conceptual design, from a designer perspective; comparisons are also made of the time and expertise required to build the geometry representations as well.
Flight control systems research. [optimization of F-8 aircraft control system
NASA Technical Reports Server (NTRS)
Whitaker, H. P.; Baram, Y.; Cheng, Y.
1973-01-01
Theoretical development is reported for the parameter optimization design technique needed for digital flight control system design. The results of an example case study applying the optimization technique for continuous systems to an F-8 aircraft feedback control system are presented. The concept of evolving the simplest system configuration that is capable of meeting a specified set of performance requirements is illustrated in this work.
Propeller design by optimization
NASA Technical Reports Server (NTRS)
Rizk, M. H.; Jou, W.-H.
1986-01-01
The feasibility of designing propellers by an optimization procedure is investigated. A scheme, which solves the full potential flow equation about a propeller by line relaxation, is modified so that the iterative solutions of the flow equation and the design parameters are updated simultaneously. Some technical problems in using optimization for designing propellers with maximum efficiency are identified. Approaches for overcoming these problems are presented.
Neighboring Optimal Aircraft Guidance in a General Wind Environment
NASA Technical Reports Server (NTRS)
Jardin, Matthew R. (Inventor)
2003-01-01
Method and system for determining an optimal route for an aircraft moving between first and second waypoints in a general wind environment. A selected first wind environment is analyzed for which a nominal solution can be determined. A second wind environment is then incorporated; and a neighboring optimal control (NOC) analysis is performed to estimate an optimal route for the second wind environment. In particular examples with flight distances of 2500 and 6000 nautical miles in the presence of constant or piecewise linearly varying winds, the difference in flight time between a nominal solution and an optimal solution is 3.4 to 5 percent. Constant or variable winds and aircraft speeds can be used. Updated second wind environment information can be provided and used to obtain an updated optimal route.
Modeling, design and energy management of fuel cell systems for aircraft
NASA Astrophysics Data System (ADS)
Bradley, Thomas Heenan
Fuel cell powered aircraft have been of long term interest to the aviation community because of their potential for improved performance and environmental compatibility. Only recently have improvements in the technological readiness of fuel cell powerplants enabled the first aviation applications of fuel cell technology. Based on the results of conceptual design studies and a few technology demonstration projects, there has emerged a widespread understanding of the importance of fuel cell powerplants for near-term and future aviation applications. Despite this, many aspects of the performance, design and construction of robust and optimized fuel cell powered aircraft have not been fully explored. This goal of this research then is to develop an improved understanding of the performance, design characteristics, design tradeoffs and viability of fuel cell powerplants for aviation applications. To accomplish these goals, new modeling, design, and experimental tools are developed, validated and applied to the design of fuel cell powered unmanned aerial vehicles. First, a general sub-system model of fuel cell powerplant performance, mass and geometry is derived from experimental and theoretical investigations of a fuel cell powerplant that is developed in hardware. These validated fuel cell subsystem models are then incorporated into a computer-based, application-integrated, parametric, and optimizeable design environment that allows for the concurrent design of the aircraft and fuel cell powerplant. The advanced modeling and design techniques required for modern aircraft design (including multi-disciplinary analysis, performance optimization under uncertainty and system performance validation), are applied at the fuel cell subsystem level and are linked to aircraft performance and design metrics. These tools and methods are then applied to the analysis and design of fuel cell powered aircraft in a series of case studies and design experiments. Based on the results of
A trade-off analysis design tool. Aircraft interior noise-motion/passenger satisfaction model
NASA Technical Reports Server (NTRS)
Jacobson, I. D.
1977-01-01
A design tool was developed to enhance aircraft passenger satisfaction. The effect of aircraft interior motion and noise on passenger comfort and satisfaction was modelled. Effects of individual aircraft noise sources were accounted for, and the impact of noise on passenger activities and noise levels to safeguard passenger hearing were investigated. The motion noise effect models provide a means for tradeoff analyses between noise and motion variables, and also provide a framework for optimizing noise reduction among noise sources. Data for the models were collected onboard commercial aircraft flights and specially scheduled tests.
Cross-Polar Aircraft Trajectory Optimization and Potential Climate Impact
NASA Technical Reports Server (NTRS)
Sridhar, Banavar; Chen, Neil; Ng, Hok
2011-01-01
Cross-Polar routes offer new opportunities for air travel markets. Transpolar flights reduce travel times, fuel burns, and associated environmental emissions by flying direct paths between many North American and Asian cities. This study evaluates the potential benefits of flying wind-optimal polar routes and assessed their potential impact on climate change. An optimization algorithm is developed for transpolar flights to generate wind-optimal trajectories that minimize climate impact of aircraft, in terms of global warming potentials (relative to warming by one kg of CO2) of several types of emissions, while avoiding regions of airspace that facilitate persistent contrail formation. Estimations of global warming potential are incorporated into the objective function of the optimization algorithm to assess the climate impact of aircraft emissions discharged at a given location and altitude. The regions of airspace with very low ambient temperature and areas favorable to persistent contrail formation are modeled as undesirable regions that aircraft should avoid and are formulated as soft state constraints. The fuel burn and climate impact of cross-polar air traffic flying various types of trajectory including flightplan, great circle, wind-optimal, and contrail-avoidance are computed for 15 origin-destination pairs between major international airports in the U.S. and Asia. Wind-optimal routes reduce average fuel burn of flight plan routes by 4.4% on December 4, 2010 and 8.0% on August 7, 2010, respectively. The tradeoff between persistent contrail formation and additional global warming potential of aircraft emissions is investigated with and without altitude optimization. Without altitude optimization, the reduction in contrail travel times is gradual with increase in total fuel consumption. When altitude is optimized, a one percent increase in additional global warming potential, a climate impact equivalent to that of 4070kg and 4220kg CO2 emission, reduces 135
Multidisciplinary Design and Analysis for Commercial Aircraft
NASA Technical Reports Server (NTRS)
Cummings, Russell M.; Freeman, H. JoAnne
1999-01-01
Multidisciplinary design and analysis (MDA) has become the normal mode of operation within most aerospace companies, but the impact of these changes have largely not been reflected at many universities. On an effort to determine if the emergence of multidisciplinary design concepts should influence engineering curricula, NASA has asked several universities (Virginia Tech, Georgia Tech, Clemson, BYU, and Cal Poly) to investigate the practicality of introducing MDA concepts within their undergraduate curricula. A multidisciplinary team of faculty, students, and industry partners evaluated the aeronautical engineering curriculum at Cal Poly. A variety of ways were found to introduce MDA themes into the curriculum without adding courses or units to the existing program. Both analytic and educational tools for multidisciplinary design of aircraft have been developed and implemented.
Structural Optimization Methodology for Rotating Disks of Aircraft Engines
NASA Technical Reports Server (NTRS)
Armand, Sasan C.
1995-01-01
In support of the preliminary evaluation of various engine technologies, a methodology has been developed for structurally designing the rotating disks of an aircraft engine. The structural design methodology, along with a previously derived methodology for predicting low-cycle fatigue life, was implemented in a computer program. An interface computer program was also developed that gathers the required data from a flowpath analysis program (WATE) being used at NASA Lewis. The computer program developed for this study requires minimum interaction with the user, thus allowing engineers with varying backgrounds in aeropropulsion to successfully execute it. The stress analysis portion of the methodology and the computer program were verified by employing the finite element analysis method. The 10th- stage, high-pressure-compressor disk of the Energy Efficient Engine Program (E3) engine was used to verify the stress analysis; the differences between the stresses and displacements obtained from the computer program developed for this study and from the finite element analysis were all below 3 percent for the problem solved. The computer program developed for this study was employed to structurally optimize the rotating disks of the E3 high-pressure compressor. The rotating disks designed by the computer program in this study were approximately 26 percent lighter than calculated from the E3 drawings. The methodology is presented herein.
Conceptual design of hybrid-electric transport aircraft
NASA Astrophysics Data System (ADS)
Pornet, C.; Isikveren, A. T.
2015-11-01
The European Flightpath 2050 and corresponding Strategic Research and Innovation Agenda (SRIA) as well as the NASA Environmentally Responsible Aviation N+ series have elaborated aggressive emissions and external noise reduction targets according to chronological waypoints. In order to deliver ultra-low or even zero in-flight emissions levels, there exists an increasing amount of international research and development emphasis on electrification of the propulsion and power systems of aircraft. Since the late 1990s, a series of experimental and a host of burgeouning commercial activities for fixed-wing aviation have focused on glider, ultra-light and light-sport airplane, and this is proving to serve as a cornerstone for more ambitious transport aircraft design and integration technical approaches. The introduction of hybrid-electric technology has dramatically expanded the design space and the full-potential of these technologies will be drawn through synergetic, tightly-coupled morphological and systems integration emphasizing propulsion - as exemplified by the potential afforded by distributed propulsion solutions. With the aim of expanding upon the current repository of knowledge associated with hybrid-electric propulsion systems a quad-fan arranged narrow-body transport aircraft equipped with two advanced Geared-Turbofans (GTF) and two Electrical Fans (EF) in an under-wing podded installation is presented in this technical article. The assessment and implications of an increasing Degree-of-Hybridization for Useful Power (HP,USE) on the overall sizing, performance as well as flight technique optimization of fuel-battery hybrid-electric aircraft is addressed herein. The integrated performance of the concept was analyzed in terms of potential block fuel burn reduction and change in vehicular efficiency in comparison to a suitably projected conventional aircraft employing GTF-only propulsion targeting year 2035. Results showed that by increasing HP,USE, significant
Adaptive Optimization of Aircraft Engine Performance Using Neural Networks
NASA Technical Reports Server (NTRS)
Simon, Donald L.; Long, Theresa W.
1995-01-01
Preliminary results are presented on the development of an adaptive neural network based control algorithm to enhance aircraft engine performance. This work builds upon a previous National Aeronautics and Space Administration (NASA) effort known as Performance Seeking Control (PSC). PSC is an adaptive control algorithm which contains a model of the aircraft's propulsion system which is updated on-line to match the operation of the aircraft's actual propulsion system. Information from the on-line model is used to adapt the control system during flight to allow optimal operation of the aircraft's propulsion system (inlet, engine, and nozzle) to improve aircraft engine performance without compromising reliability or operability. Performance Seeking Control has been shown to yield reductions in fuel flow, increases in thrust, and reductions in engine fan turbine inlet temperature. The neural network based adaptive control, like PSC, will contain a model of the propulsion system which will be used to calculate optimal control commands on-line. Hopes are that it will be able to provide some additional benefits above and beyond those of PSC. The PSC algorithm is computationally intensive, it is valid only at near steady-state flight conditions, and it has no way to adapt or learn on-line. These issues are being addressed in the development of the optimal neural controller. Specialized neural network processing hardware is being developed to run the software, the algorithm will be valid at steady-state and transient conditions, and will take advantage of the on-line learning capability of neural networks. Future plans include testing the neural network software and hardware prototype against an aircraft engine simulation. In this paper, the proposed neural network software and hardware is described and preliminary neural network training results are presented.
Design of a turbofan powered regional transport aircraft
NASA Technical Reports Server (NTRS)
1991-01-01
The majority of the market for small commercial transport aircraft is dominated by high efficiency propeller driven aircraft of non-U.S. manufacture. During the past year, an aircraft was designed with ranges of up to 1500 nautical miles and passenger loads between 50 and 90. Special emphasis was placed upon keeping acquisition cost and direct operating costs at a low level while providing passengers with quality comfort levels. Several designs are presented which place a high premium on design innovation.
Aircraft Range Optimization Using Singular Perturbations
NASA Technical Reports Server (NTRS)
Oconnor, Joseph Taffe
1973-01-01
An approximate analytic solution is developed for the problem of maximizing the range of an aircraft for a fixed end state. The problem is formulated as a singular perturbation and solved by matched inner and outer asymptotic expansions and the minimum principle of Pontryagin. Cruise in the stratosphere, and on transition to and from cruise at constant Mach number are discussed. The state vector includes altitude, flight path angle, and mass. Specific fuel consumption becomes a linear function of power approximating that of the cruise values. Cruise represents the outer solution; altitude and flight path angle are constants, and only mass changes. Transitions between cruise and the specified initial and final conditions correspond to the inner solutions. The mass is constant and altitude and velocity vary. A solution is developed which is valid for cruise but which is not for the initial and final conditions. Transforming of the independent variable near the initial and final conditions result in solutions which are valid for the two inner solutions but not for cruise. The inner solutions can not be obtained without simplifying the state equations. The singular perturbation approach overcomes this difficulty. A quadratic approximation of the state equations is made. The resulting problem is solved analytically, and the two inner solutions are matched to the outer solution.
Optimizing aircraft performance with adaptive, integrated flight/propulsion control
NASA Technical Reports Server (NTRS)
Smith, R. H.; Chisholm, J. D.; Stewart, J. F.
1991-01-01
The Performance-Seeking Control (PSC) integrated flight/propulsion adaptive control algorithm presented was developed in order to optimize total aircraft performance during steady-state engine operation. The PSC multimode algorithm minimizes fuel consumption at cruise conditions, while maximizing excess thrust during aircraft accelerations, climbs, and dashes, and simultaneously extending engine service life through reduction of fan-driving turbine inlet temperature upon engagement of the extended-life mode. The engine models incorporated by the PSC are continually upgraded, using a Kalman filter to detect anomalous operations. The PSC algorithm will be flight-demonstrated by an F-15 at NASA-Dryden.
NASA Technical Reports Server (NTRS)
Sandlin, Doral R.; Bauer, Brent Alan
1993-01-01
This paper discusses the development of a FORTRAN computer code to perform agility analysis on aircraft configurations. This code is to be part of the NASA-Ames ACSYNT (AirCraft SYNThesis) design code. This paper begins with a discussion of contemporary agility research in the aircraft industry and a survey of a few agility metrics. The methodology, techniques and models developed for the code are then presented. Finally, example trade studies using the agility module along with ACSYNT are illustrated. These trade studies were conducted using a Northrop F-20 Tigershark aircraft model. The studies show that the agility module is effective in analyzing the influence of common parameters such as thrust-to-weight ratio and wing loading on agility criteria. The module can compare the agility potential between different configurations. In addition one study illustrates the module's ability to optimize a configuration's agility performance.
Optimization of digital designs
NASA Technical Reports Server (NTRS)
Whitaker, Sterling R. (Inventor); Miles, Lowell H. (Inventor)
2009-01-01
An application specific integrated circuit is optimized by translating a first representation of its digital design to a second representation. The second representation includes multiple syntactic expressions that admit a representation of a higher-order function of base Boolean values. The syntactic expressions are manipulated to form a third representation of the digital design.
Linear tracking systems with applications to aircraft control system design
NASA Technical Reports Server (NTRS)
Lee, W. H.; Athans, M.; Castanon, D.; Bacchioloni, F.
1977-01-01
A class of optimal linear time invariant tracking systems, both in continuous time and discrete time, of which the number of inputs (which are restricted to be step functions) is equal to the number of system outputs, is studied. Along with derivation of equations and design procedures, two discretization schemes are presented, constraining either the control or its time derivative, to be a constant over each sampling period. Descriptions are given for the linearized model of the F-8C aircraft longitudinal dynamics, and the C* handling qualities criterion, which then serve as an illustration of the applications of these linear tracking designs. A suboptimal reduced state design is also presented. Numerical results are given for both the continuous time and discrete time designs.
Optimization via CFD of aircraft hot-air anti-icing systems
NASA Astrophysics Data System (ADS)
Pellissier, Mathieu Paul Constantin
In-flight icing is a major concern in aircraft safety and a non-negligible source of incidents and accidents, and is still a serious hazard today. It remains consequently a design and certification challenge for aircraft manufacturers. The aerodynamic performance of an aircraft can indeed degrade rapidly when flying in icing conditions, leading to incidents or accidents. In-flight icing occurs when an aircraft passes through clouds containing supercooled water droplets at or below freezing temperature. Droplets impinge on its exposed surfaces and freeze, causing roughness and shape changes that increase drag, decrease lift and reduce the stall angle of attack, eventually inducing flow separation and stall. This hazardous ice accretion is prevented by the use of dedicated anti-icing systems, among which hot-air-types are the most common for turbofan aircraft. This work presents a methodology for the optimization of such aircraft hot-air-type anti-icing systems, known as Piccolo tubes. Having identified through 3D Computational Fluid Dynamics (CFD) the most critical in-flight icing conditions, as well as determined thermal power constraints, the objective is to optimize the heat distribution in such a way to minimize power requirements, while meeting or exceeding all safety regulation requirements. To accomplish this, an optimization method combining 3D CFD, Reduced-Order Models (ROM) and Genetic Algorithms (GA) is constructed to determine the optimal configuration of the Piccolo tube (angles of jets, spacing between holes, and position from leading edge). The methodology successfully results in increasingly optimal configurations from three up to five design variables.
Landing Gear Integration in Aircraft Conceptual Design. Revision
NASA Technical Reports Server (NTRS)
Chai, Sonny T.; Mason, William H.
1997-01-01
The design of the landing gear is one of the more fundamental aspects of aircraft design. The design and integration process encompasses numerous engineering disciplines, e.g., structure, weights, runway design, and economics, and has become extremely sophisticated in the last few decades. Although the design process is well-documented, no attempt has been made until now in the development of a design methodology that can be used within an automated environment. As a result, the process remains to be a key responsibility for the configuration designer and is largely experience-based and graphically-oriented. However, as industry and government try to incorporate multidisciplinary design optimization (MDO) methods in the conceptual design phase, the need for a more systematic procedure has become apparent. The development of an MDO-capable design methodology as described in this work is focused on providing the conceptual designer with tools to help automate the disciplinary analyses, i.e., geometry, kinematics, flotation, and weight. Documented design procedures and analyses were examined to determine their applicability, and to ensure compliance with current practices and regulations. Using the latest information as obtained from industry during initial industry survey, the analyses were in terms modified and expanded to accommodate the design criteria associated with the advanced large subsonic transports. Algorithms were then developed based on the updated analysis procedures to be incorporated into existing MDO codes.
Aircraft path planning for optimal imaging using dynamic cost functions
NASA Astrophysics Data System (ADS)
Christie, Gordon; Chaudhry, Haseeb; Kochersberger, Kevin
2015-05-01
Unmanned aircraft development has accelerated with recent technological improvements in sensing and communications, which has resulted in an "applications lag" for how these aircraft can best be utilized. The aircraft are becoming smaller, more maneuverable and have longer endurance to perform sensing and sampling missions, but operating them aggressively to exploit these capabilities has not been a primary focus in unmanned systems development. This paper addresses a means of aerial vehicle path planning to provide a realistic optimal path in acquiring imagery for structure from motion (SfM) reconstructions and performing radiation surveys. This method will allow SfM reconstructions to occur accurately and with minimal flight time so that the reconstructions can be executed efficiently. An assumption is made that we have 3D point cloud data available prior to the flight. A discrete set of scan lines are proposed for the given area that are scored based on visibility of the scene. Our approach finds a time-efficient path and calculates trajectories between scan lines and over obstacles encountered along those scan lines. Aircraft dynamics are incorporated into the path planning algorithm as dynamic cost functions to create optimal imaging paths in minimum time. Simulations of the path planning algorithm are shown for an urban environment. We also present our approach for image-based terrain mapping, which is able to efficiently perform a 3D reconstruction of a large area without the use of GPS data.
Analysis of LPFG sensor systems for aircraft wing drag optimization
NASA Astrophysics Data System (ADS)
Kazemi, Alex A.; Ishihara, Abe
2014-09-01
In normal fiber, the refractive indices of the core and cladding do not change along the length of the fiber; however, by inducing a periodic modulation of refractive index along the length in the core of the optical fiber, the optical fiber grating is produced. This exhibits very interesting spectral properties and for this reason we propose to develop and integrate a distributed sensor network based on long period fiber gratings (LPFGs) technology which has grating periods on the order of 100 Î¼m to 1 mm to be embedded in the wing section of aircraft to measure bending and torsion in real-time in order to measure wing deformation of commercial airplanes resulting in extensive benefits such as reduced structural weight, mitigation of induced drag and lower fuel consumption which is fifty percent of total cost of operation for airline industry. Fiber optic sensors measurement capabilities are as vital as they are for other sensing technologies, but optical measurements differ in important ways. In this paper we focus on the testing and aviation requirements for LPFG sensors. We discuss the bases of aviation standards for fiber optic sensor measurements, and the quantities that are measured. Our main objective is to optimize the design for material, mechanical, optical and environmental requirements. We discuss the analysis and evaluation of extensive testing of LPFG sensor systems such as attenuation, environmental, humidity, fluid immersion, temperature cycling, aging, smoke, flammability, impact resistance, flexure endurance, tensile, vitiation and shock.
Application of optimal input synthesis to aircraft parameter identification
NASA Technical Reports Server (NTRS)
Gupta, N. K.; Hall, W. E., Jr.; Mehra, R. K.
1976-01-01
The Frequency Domain Input Synthesis procedure is used in identifying the stability and control derivatives of an aircraft. By using a frequency-domain approach, one can handle criteria that are not easily handled by the time-domain approaches. Numerical results are presented for optimal elevator deflections to estimate the longitudinal stability and control derivatives subject to root-mean square constraints on the input. The applicability of the steady state optimal inputs to finite duration flight testing is investigated. The steady state approximation of frequency-domain synthesis is good for data lengths greater than two time cycles for the short period mode of the aircraft longitudinal motions. Phase relationships between different frequency components become important for shorter data lengths. The frequency domain inputs are shown to be much better than the conventional doublet inputs.
Simulation Tools Model Icing for Aircraft Design
NASA Technical Reports Server (NTRS)
2012-01-01
Here s a simple science experiment to try: Place an unopened bottle of distilled water in your freezer. After 2-3 hours, if the water is pure enough, you will notice that it has not frozen. Carefully pour the water into a bowl with a piece of ice in it. When it strikes the ice, the water will instantly freeze. One of the most basic and commonly known scientific facts is that water freezes at around 32 F. But this is not always the case. Water lacking any impurities for ice crystals to form around can be supercooled to even lower temperatures without freezing. High in the atmosphere, water droplets can achieve this delicate, supercooled state. When a plane flies through clouds containing these droplets, the water can strike the airframe and, like the supercooled water hitting the ice in the experiment above, freeze instantly. The ice buildup alters the aerodynamics of the plane - reducing lift and increasing drag - affecting its performance and presenting a safety issue if the plane can no longer fly effectively. In certain circumstances, ice can form inside aircraft engines, another potential hazard. NASA has long studied ways of detecting and countering atmospheric icing conditions as part of the Agency s efforts to enhance aviation safety. To do this, the Icing Branch at Glenn Research Center utilizes a number of world-class tools, including the Center s Icing Research Tunnel and the NASA 607 icing research aircraft, a "flying laboratory" for studying icing conditions. The branch has also developed a suite of software programs to help aircraft and icing protection system designers understand the behavior of ice accumulation on various surfaces and in various conditions. One of these innovations is the LEWICE ice accretion simulation software. Initially developed in the 1980s (when Glenn was known as Lewis Research Center), LEWICE has become one of the most widely used tools in icing research and aircraft design and certification. LEWICE has been transformed over
Aerodynamic Design Opportunities for Future Supersonic Aircraft
NASA Technical Reports Server (NTRS)
Wood, Richard M.; Bauer, Steven X. S.; Flamm, Jeffrey D.
2002-01-01
A discussion of a diverse set of aerodynamic opportunities to improve the aerodynamic performance of future supersonic aircraft has been presented and discussed. These ideas are offered to the community in a hope that future supersonic vehicle development activities will not be hindered by past efforts. A number of nonlinear flow based drag reduction technologies are presented and discussed. The subject technologies are related to the areas of interference flows, vehicle concepts, vortex flows, wing design, advanced control effectors, and planform design. The authors also discussed the importance of improving the aerodynamic design environment to allow creativity and knowledge greater influence. A review of all of the data presented show that pressure drag reductions on the order of 50 to 60 counts are achievable, compared to a conventional supersonic cruise vehicle, with the application of several of the discussed technologies. These drag reductions would correlate to a 30 to 40% increase in cruise L/D (lift-to-drag ratio) for a commercial supersonic transport.
Multidisciplinary design optimization for sonic boom mitigation
NASA Astrophysics Data System (ADS)
Ozcer, Isik A.
product design. The simulation tools are used to optimize three geometries for sonic boom mitigation. The first is a simple axisymmetric shape to be used as a generic nose component, the second is a delta wing with lift, and the third is a real aircraft with nose and wing optimization. The objectives are to minimize the pressure impulse or the peak pressure in the sonic boom signal, while keeping the drag penalty under feasible limits. The design parameters for the meridian profile of the nose shape are the lengths and the half-cone angles of the linear segments that make up the profile. The design parameters for the lifting wing are the dihedral angle, angle of attack, non-linear span-wise twist and camber distribution. The test-bed aircraft is the modified F-5E aircraft built by Northrop Grumman, designated the Shaped Sonic Boom Demonstrator. This aircraft is fitted with an optimized axisymmetric nose, and the wings are optimized to demonstrate optimization for sonic boom mitigation for a real aircraft. The final results predict 42% reduction in bow shock strength, 17% reduction in peak Deltap, 22% reduction in pressure impulse, 10% reduction in foot print size, 24% reduction in inviscid drag, and no loss in lift for the optimized aircraft. Optimization is carried out using response surface methodology, and the design matrices are determined using standard DoE techniques for quadratic response modeling.
Optimization of aircraft seat cushion fire blocking layers
NASA Technical Reports Server (NTRS)
Kourtides, D. A.; Parker, J. A.; Ling, A. C.; Hovatter, W. R.
1983-01-01
This report describes work completed by the National Aeronautics and Space Administration - for the Federal Aviation Administration Technical Center. The purpose of this work was to examine the potential of fire blocking mechanisms for aircraft seat cushions in order to provide an optimized seat configuration with adequate fire protection and minimum weight. Aluminized thermally stable fabrics were found to provide adequate fire protection when used in conjunction with urethane foams, while maintaining minimum weight and cost penalty.
Optimal Discrete Event Supervisory Control of Aircraft Gas Turbine Engines
NASA Technical Reports Server (NTRS)
Litt, Jonathan (Technical Monitor); Ray, Asok
2004-01-01
This report presents an application of the recently developed theory of optimal Discrete Event Supervisory (DES) control that is based on a signed real measure of regular languages. The DES control techniques are validated on an aircraft gas turbine engine simulation test bed. The test bed is implemented on a networked computer system in which two computers operate in the client-server mode. Several DES controllers have been tested for engine performance and reliability.
Some comparisons of US and USSR aircraft design developments
NASA Technical Reports Server (NTRS)
Spearman, M. L.
1985-01-01
A review is given of the design and development of some U.S. and U.S.S.R. aircraft. The emphasis is on the historical development of large aircraft - civil and military transports and bombers. Design trends are somewhat similar for the two countries and indications are that some fundamental characteristics are dictated more by ideological differences rather than technological differences. A brief description is given in a more or less chronological order of the major bomber aircraft, major civil and military transport aircraft, and the development of the air transport systems.
Some comparisons of US and USSR aircraft design developments
NASA Technical Reports Server (NTRS)
Spearman, M. L.
1985-01-01
A review is given of the design and development of some US and USSR aircraft. The emphasis is on the historical development of large aircraft-civil and military transports and bombers. Design trends are somewhat similar for the two countries and indications are that some fundamental characteristics are dictated more by ideological differences rather than technological differences. A brief description is given in a more or less chronological order of the major bomber aircraft, major civil and military transport aircraft, and the development of the air transport systems.
The design of a long range megatransport aircraft
NASA Technical Reports Server (NTRS)
Weisshaar, Terrence A.; Allen, Carl L.
1992-01-01
During the period from August 1991 - June 1992 two design classes at Purdue University participated in the design of a long range, high capacity transport aircraft, dubbed the megatransport. Thirteen Purdue design teams generated RFP's that defined passenger capability and range, based upon team perception of market needs and infrastructure constraints. Turbofan engines were designed by each group to power these aircraft. The design problem and the variety of solutions developed are described in an attached paper.
Conceptual design of single turbofan engine powered light aircraft
NASA Technical Reports Server (NTRS)
Snyder, F. S.; Voorhees, C. G.; Heinrich, A. M.; Baisden, D. N.
1977-01-01
The conceptual design of a four place single turbofan engine powered light aircraft was accomplished utilizing contemporary light aircraft conventional design techniques as a means of evaluating the NASA-Ames General Aviation Synthesis Program (GASP) as a preliminary design tool. In certain areas, disagreement or exclusion were found to exist between the results of the conventional design and GASP processes. Detail discussion of these points along with the associated contemporary design methodology are presented.
Use of optimization to predict the effect of selected parameters on commuter aircraft performance
NASA Technical Reports Server (NTRS)
Wells, V. L.; Shevell, R. S.
1982-01-01
The relationships between field length and cruise speed and aircraft direct operating cost were determined. A gradient optimizing computer program was developed to minimize direct operating cost (DOC) as a function of airplane geometry. In this way, the best airplane operating under one set of constraints can be compared with the best operating under another. A constant 30-passenger fuselage and rubberized engines based on the General Electric CT-7 were used as a baseline. All aircraft had to have a 600 nautical mile maximum range and were designed to FAR part 25 structural integrity and climb gradient regulations. Direct operating cost was minimized for a typical design mission of 150 nautical miles. For purposes of C sub L sub max calculation, all aircraft had double-slotted flaps but with no Fowler action.
Figures of merit for airfoil/aircraft design integration
NASA Technical Reports Server (NTRS)
Maughmer, Mark D.; Somers, Dan M.
1988-01-01
Because the airfoil can so strongly impact other aspects of an aircraft configuration, it is important that the airfoil design process be integrated with that of the aircraft to achieve the best possible performance of a new flight vehicle. To aid in preliminary design efforts, several aerodynamic figures of merit are presented which facilitate the matching of the airfoil performance characteristics to those of the aircraft. These figures of merit are fairly general and can assist the airfoil design process for flight vehicles designed for maximum endurance, range, or ceiling. Although specifically applicable to vehicles for which the wing area is sized by some required minimum airspeed, the discussion is pertinent to all airfoil/aircraft matching situations and points the way for developing similar figures of merit to aid the airfoil/aircraft design process for any flight vehicle.
Conceptual design of high speed supersonic aircraft: A brief review on SR-71 (Blackbird) aircraft
NASA Astrophysics Data System (ADS)
Xue, Hui; Khawaja, H.; Moatamedi, M.
2014-12-01
The paper presents the conceptual design of high-speed supersonic aircraft. The study focuses on SR-71 (Blackbird) aircraft. The input to the conceptual design is a mission profile. Mission profile is a flight profile of the aircraft defined by the customer. This paper gives the SR-71 aircraft mission profile specified by US air force. Mission profile helps in defining the attributes the aircraft such as wing profile, vertical tail configuration, propulsion system, etc. Wing profile and vertical tail configurations have direct impact on lift, drag, stability, performance and maneuverability of the aircraft. A propulsion system directly influences the performance of the aircraft. By combining the wing profile and the propulsion system, two important parameters, known as wing loading and thrust to weight ratio can be calculated. In this work, conceptual design procedure given by D. P. Raymer (AIAA Educational Series) is applied to calculate wing loading and thrust to weight ratio. The calculated values are compared against the actual values of the SR-71 aircraft. Results indicates that the values are in agreement with the trend of developments in aviation.
Parallel calculation of sensitivity derivatives for aircraft design using automatic differentiation
Bischof, C.H.; Knauff, T.L. Jr.; Green, L.L.; Haigler, K.J.
1994-01-01
Realistic multidisciplinary design optimization (MDO) of advanced aircraft using state-of-the-art computers is an extremely challenging problem from both the physical modelling and computer science points of view. In order to produce an efficient aircraft design, many trade-offs must be made among the various physical design variables. Similarly, in order to produce an efficient design scheme, many trade-offs must be made among the various MDO implementation options. In this paper, we examine the effects of vectorization and coarse-grained parallelization on the SD calculation using a representative example taken from a transonic transport design problem.
A methodology for designing aircraft to low sonic boom constraints
NASA Technical Reports Server (NTRS)
Mack, Robert J.; Needleman, Kathy E.
1991-01-01
A method for designing conceptual supersonic cruise aircraft to meet low sonic boom requirements is outlined and described. The aircraft design is guided through a systematic evolution from initial three view drawing to a final numerical model description, while the designer using the method controls the integration of low sonic boom, high supersonic aerodynamic efficiency, adequate low speed handling, and reasonable structure and materials technologies. Some experience in preliminary aircraft design and in the use of various analytical and numerical codes is required for integrating the volume and lift requirements throughout the design process.
Conceptual design of an aircraft automated coating removal system
Baker, J.E.; Draper, J.V.; Pin, F.G.; Primm, A.H.; Shekhar, S.
1996-05-01
Paint stripping of the U.S. Air Force`s large transport aircrafts is currently a labor-intensive, manual process. Significant reductions in costs, personnel and turnaround time can be accomplished by the judicious use of automation in some process tasks. This paper presents the conceptual design of a coating removal systems for the tail surfaces of the C-5 plane. Emphasis is placed on the technology selection to optimize human-automation synergy with respect to overall costs, throughput, quality, safety, and reliability. Trade- offs between field-proven vs. research-requiring technologies, and between expected gain vs. cost and complexity, have led to a conceptual design which is semi-autonomous (relying on the human for task specification and disturbance handling) yet incorporates sensor- based automation (for sweep path generation and tracking, surface following, stripping quality control and tape/breach handling).
The design of a long-range megatransport aircraft
NASA Technical Reports Server (NTRS)
Weisshaar, Terrence A.; Allen, Carl L.
1992-01-01
Aircraft manufacturers are examining the market and feasibility of long-range passenger aircraft carrying more than 600 passengers. These aircraft would carry travelers at reduced cost and, at the same time, reduce congestion around major airports. The design of a large, long-range transport involves broad issues such as: the integration of airport terminal facilities; passenger loading and unloading; trade-offs between aircraft size and the cost to reconfigure these existing facilities; and, defeating the 'square-cube' law. Thirteen Purdue design teams generated RFP's that defined passenger capability and range, based upon team perception of market needs and infrastructure constraints. Turbofan engines were designed by each group to power these aircraft. The design problem and the variety of solutions developed are reviewed.
MDO can help resolve the designer's dilemma. [Multidisciplinary design optimization
Sobieszczanski-sobieski, Jaroslaw; Tulinius, J.R. Rockwell International Corp., El Segundo, CA )
1991-09-01
Multidisciplinary design optimization (MDO) is presented as a rapidly growing body of methods, algorithms, and techniques that will provide a quantum jump in the effectiveness and efficiency of the quantitative side of design, and will turn that side into an environment in which the qualitative side can thrive. MDO borrows from CAD/CAM for graphic visualization of geometrical and numerical data, data base technology, and in computer software and hardware. Expected benefits from this methodology are a rational, mathematically consistent approach to hypersonic aircraft designs, designs pushed closer to the optimum, and a design process either shortened or leaving time available for different concepts to be explored.
MDO can help resolve the designer's dilemma. [multidisciplinary design optimization
NASA Technical Reports Server (NTRS)
Sobieszczanski-Sobieski, Jaroslaw; Tulinius, Jan R.
1991-01-01
Multidisciplinary design optimization (MDO) is presented as a rapidly growing body of methods, algorithms, and techniques that will provide a quantum jump in the effectiveness and efficiency of the quantitative side of design, and will turn that side into an environment in which the qualitative side can thrive. MDO borrows from CAD/CAM for graphic visualization of geometrical and numerical data, data base technology, and in computer software and hardware. Expected benefits from this methodology are a rational, mathematically consistent approach to hypersonic aircraft designs, designs pushed closer to the optimum, and a design process either shortened or leaving time available for different concepts to be explored.
The role of computational fluid dynamics (CFD) in aircraft design
Tinoco, E.N. )
1990-01-01
The application of CFD to aircraft design configurations and its influence on the aircraft development and support process is analyzed. Results indicate that combining CFD and the wind tunnel can achieve design solutions that otherwise would not be found, and can also significantly reduce the length of the design cycle. It is concluded that CFD provides for a better understanding of flow physics, achievement of design solutions that are otherwise unobtainable, and reduction of development flowtime.
NASA Technical Reports Server (NTRS)
Sensmeier, Mark D.; Samareh, Jamshid A.
2005-01-01
An approach is proposed for the application of rapid generation of moderate-fidelity structural finite element models of air vehicle structures to allow more accurate weight estimation earlier in the vehicle design process. This should help to rapidly assess many structural layouts before the start of the preliminary design phase and eliminate weight penalties imposed when actual structure weights exceed those estimated during conceptual design. By defining the structural topology in a fully parametric manner, the structure can be mapped to arbitrary vehicle configurations being considered during conceptual design optimization. A demonstration of this process is shown for two sample aircraft wing designs.
Multidisciplinary design optimization using multiobjective formulation techniques
NASA Technical Reports Server (NTRS)
Chattopadhyay, Aditi; Pagaldipti, Narayanan S.
1995-01-01
This report addresses the development of a multidisciplinary optimization procedure using an efficient semi-analytical sensitivity analysis technique and multilevel decomposition for the design of aerospace vehicles. A semi-analytical sensitivity analysis procedure is developed for calculating computational grid sensitivities and aerodynamic design sensitivities. Accuracy and efficiency of the sensitivity analysis procedure is established through comparison of the results with those obtained using a finite difference technique. The developed sensitivity analysis technique are then used within a multidisciplinary optimization procedure for designing aerospace vehicles. The optimization problem, with the integration of aerodynamics and structures, is decomposed into two levels. Optimization is performed for improved aerodynamic performance at the first level and improved structural performance at the second level. Aerodynamic analysis is performed by solving the three-dimensional parabolized Navier Stokes equations. A nonlinear programming technique and an approximate analysis procedure are used for optimization. The proceduredeveloped is applied to design the wing of a high speed aircraft. Results obtained show significant improvements in the aircraft aerodynamic and structural performance when compared to a reference or baseline configuration. The use of the semi-analytical sensitivity technique provides significant computational savings.
Performance Evaluation Method for Dissimilar Aircraft Designs
NASA Technical Reports Server (NTRS)
Walker, H. J.
1979-01-01
A rationale is presented for using the square of the wingspan rather than the wing reference area as a basis for nondimensional comparisons of the aerodynamic and performance characteristics of aircraft that differ substantially in planform and loading. Working relationships are developed and illustrated through application to several categories of aircraft covering a range of Mach numbers from 0.60 to 2.00. For each application, direct comparisons of drag polars, lift-to-drag ratios, and maneuverability are shown for both nondimensional systems. The inaccuracies that may arise in the determination of aerodynamic efficiency based on reference area are noted. Span loading is introduced independently in comparing the combined effects of loading and aerodynamic efficiency on overall performance. Performance comparisons are made for the NACA research aircraft, lifting bodies, century-series fighter aircraft, F-111A aircraft with conventional and supercritical wings, and a group of supersonic aircraft including the B-58 and XB-70 bomber aircraft. An idealized configuration is included in each category to serve as a standard for comparing overall efficiency.
NASA Technical Reports Server (NTRS)
Conlon, J. A.; Bowles, J. V.
1978-01-01
With an overall goal of defining the needs and requirements for short-haul transport aircraft research and development, the objective of this paper is to determine the performance and noise impact of short-haul transport aircraft designed with an advanced turboprop propulsion system. This propulsion system features high-speed propellers that have more blades and reduced diameters. Aircraft are designed for short and medium field lengths; mission block fuel and direct operating costs (DOC) are used as performance measures. The propeller diameter was optimized to minimize DOC. Two methods are employed to estimate the weight of the acoustic treatment needed to reduce interior noise to an acceptable level. Results show decreasing gross weight, block fuel, DOC, engine size, and optimum propfan diameter with increasing field length. The choice of acoustic treatment method has a significant effect on the aircraft design.
The optimal control frequency response problem in manual control. [of manned aircraft systems
NASA Technical Reports Server (NTRS)
Harrington, W. W.
1977-01-01
An optimal control frequency response problem is defined within the context of the optimal pilot model. The problem is designed to specify pilot model control frequencies reflective of important aircraft system properties, such as control feel system dynamics, airframe dynamics, and gust environment, as well as man machine properties, such as task and attention allocation. This is accomplished by determining a bounded set of control frequencies which minimize the total control cost. The bounds are given by zero and the neuromuscular control frequency response for each control actuator. This approach is fully adaptive, i.e., does not depend upon user entered estimates. An algorithm is developed to solve this optimal control frequency response problem. The algorithm is then applied to an attitude hold task for a bare airframe fighter aircraft case with interesting dynamic properties.
Fighter aircraft flight control technology design requirements
NASA Technical Reports Server (NTRS)
Nelson, W. E., Jr.
1984-01-01
The evolution of fighter aircraft flight control technology is briefly surveyed. Systems engineering, battle damage considerations for adaptive flutter suppression, in-flight simulation, and artificial intelligence are briefly discussed.
Optimal Aircraft Control Upset Recovery With and Without Component Failures
NASA Technical Reports Server (NTRS)
Sparks, Dean W.; Moerder, Daniel D.
2002-01-01
This paper treats the problem of recovering sustainable nondescending (safe) flight in a transport aircraft after one or more of its control effectors fail. Such recovery can be a challenging goal for many transport aircraft currently in the operational fleet for two reasons. First, they have very little redundancy in their means of generating control forces and moments. These aircraft have, as primary control surfaces, a single rudder and pairwise elevators and aileron/spoiler units that provide yaw, pitch, and roll moments with sufficient bandwidth to be used in stabilizing and maneuvering the airframe. Beyond this, throttling the engines can provide additional moments, but on a much slower time scale. Other aerodynamic surfaces, such as leading and trailing edge flaps, are only intended to be placed in a position and left, and are, hence, very slow-moving. Because of this, loss of a primary control surface strongly degrades the controllability of the vehicle, particularly when the failed effector becomes stuck in a non-neutral position where it exerts a disturbance moment that must be countered by the remaining operating effectors. The second challenge in recovering safe flight is that these vehicles are not agile, nor can they tolerate large accelerations. This is of special importance when, at the outset of the recovery maneuver, the aircraft is flying toward the ground, as is frequently the case when there are major control hardware failures. Recovery of safe flight is examined in this paper in the context of trajectory optimization. For a particular transport aircraft, and a failure scenario inspired by an historical air disaster, recovery scenarios are calculated with and without control surface failures, to bring the aircraft to safe flight from the adverse flight condition that it had assumed, apparently as a result of contact with a vortex from a larger aircraft's wake. An effort has been made to represent relevant airframe dynamics, acceleration limits
NASA Technical Reports Server (NTRS)
Hopkins, Dale A.
1998-01-01
A key challenge in designing the new High Speed Civil Transport (HSCT) aircraft is determining a good match between the airframe and engine. Multidisciplinary design optimization can be used to solve the problem by adjusting parameters of both the engine and the airframe. Earlier, an example problem was presented of an HSCT aircraft with four mixed-flow turbofan engines and a baseline mission to carry 305 passengers 5000 nautical miles at a cruise speed of Mach 2.4. The problem was solved by coupling NASA Lewis Research Center's design optimization testbed (COMETBOARDS) with NASA Langley Research Center's Flight Optimization System (FLOPS). The computing time expended in solving the problem was substantial, and the instability of the FLOPS analyzer at certain design points caused difficulties. In an attempt to alleviate both of these limitations, we explored the use of two approximation concepts in the design optimization process. The two concepts, which are based on neural network and linear regression approximation, provide the reanalysis capability and design sensitivity analysis information required for the optimization process. The HSCT aircraft optimization problem was solved by using three alternate approaches; that is, the original FLOPS analyzer and two approximate (derived) analyzers. The approximate analyzers were calibrated and used in three different ranges of the design variables; narrow (interpolated), standard, and wide (extrapolated).
Review of evolving trends in blended wing body aircraft design
NASA Astrophysics Data System (ADS)
Okonkwo, Paul; Smith, Howard
2016-04-01
The desire to produce environmentally friendly aircraft that is aerodynamically efficient and capable of conveying large number of passengers over long ranges at reduced direct operating cost led aircraft designers to develop the Blended Wing Body (BWB) aircraft concept. The BWB aircraft represents a paradigm shift in the design of aircraft. The design provides aerodynamics and environmental benefits and is suitable for the integration of advanced systems and concepts like laminar flow technology, jet flaps and distributed propulsion. However, despite these benefits, the BWB is yet to be developed for commercial air transport due to several challenges. This paper reviews emerging trends in BWB aircraft design highlighting design challenges that have hindered the development of a BWB passenger transport aircraft. The study finds that in order to harness the advantages and reduce the deficiencies of a tightly coupled configuration like the BWB, a multidisciplinary design synthesis optimisation should be conducted with good handling and ride quality as objective functions within acceptable direct operating cost and noise bounds.
Design of a turbofan powered regional transport aircraft
NASA Technical Reports Server (NTRS)
1991-01-01
The majority of the market for small commercial transport aircraft is dominated by high-efficiency, propeller-driven aircraft of non-U.S. manufacture. During the past year senior student design teams at Purdue developed and then responded to a Request For Proposal (RFP) for a regional transport aircraft. The RFP development identified promising world markets and their needs. The students responded by designing aircraft with ranges of up to 1500 n.m. and passenger loads of 50 to 90. During the design project, special emphasis was placed upon keeping acquisition cost and direct operating costs at a low level while providing passengers with quality comfort levels. Twelve student teams worked for one semester developing their designs. Several of the more successful designs and those that placed a high premium on innovation are described. The depth of detail and analysis in these student efforts are also illustrated.
The effects of aircraft design on STOL ride quality
NASA Technical Reports Server (NTRS)
Jones, C. R.; Jacobson, I. D.
1975-01-01
Effects of aircraft dynamic characteristics on passenger ride quality are investigated to determine ride-quality isocontours similar to aircraft handling-qualities contours. Measurements are made on a moving-base simulator while varying the aircraft short-period and Dutch Roll frequencies and dampings. Both pilot ratings and subjective ride-quality ratings are obtained during flight. Ride and handling qualities were found to be complementary for the Dutch Roll mode, but not for the short-period mode. Regions of optimal ride and handling qualities are defined for the short-period mode, and the effects of turbulence levels studied.
Active and passive structural design concepts for improved empennage effectiveness of aircraft
NASA Astrophysics Data System (ADS)
Weiss, Franz; Schweiger, Johannes; Simpson, John; Kullrich, Thomas
2000-06-01
Active structures concepts for the design of aircraft have been investigated for several years. Concerning static aeroelastic applications, all concepts known to the authors are trying to improve the design of aircraft wings. In the case of wings however, the design space for active structures concepts is limited by a multitude of functional requirements. Empennage surfaces on the other hand only have to meet two basic requirements: sufficient stability and maneuverability for the longitudinal and lateral motion of the aircraft. In the case of vertical tails, the aerodynamic effectiveness for the side force and for the rudder yawing moment are usually reduced by the flexibility of the structure. This causes a weight increase for the structure, which is especially unpleasant for tail surfaces because of the rearward shift of the center of gravity. Today, multidisciplinary structural optimization methods can be used to minimize the weight penalty for static aeroelastic effectiveness requirements. But an amount of penalty still remains. A smart solution for additional weight savings, if possible below the conventional basic strength design of the structure, would therefore be very welcome for any new aircraft design. The paper will describe a new design approach for vertical tails. The concept is based on a smart system for the attachment of the complete tail surface to the fuselage. If properly designed, the variable stiffness of this system will provide improved aerodynamic effectiveness of the tail at any flight condition compared to the rigid aircraft. In a first step, the structure for the vertical tail of a fighter aircraft is designed for static strength and buckling stability by means of a structural optimization program, which is based on finite element methods. The impacts of static aeroelastic effectiveness and flutter stability criteria on the structural design are shown. A modified structural model is then used to incorporate the active system for the
Numeric Design and Performance Analysis of Solid Oxide Fuel Cell -- Gas Turbine Hybrids on Aircraft
NASA Astrophysics Data System (ADS)
Hovakimyan, Gevorg
The aircraft industry benefits greatly from small improvements in aircraft component design. One possible area of improvement is in the Auxiliary Power Unit (APU). Modern aircraft APUs are gas turbines located in the tail section of the aircraft that generate additional power when needed. Unfortunately the efficiency of modern aircraft APUs is low. Solid Oxide Fuel Cell/Gas Turbine (SOFC/GT) hybrids are one possible alternative for replacing modern gas turbine APUs. This thesis investigates the feasibility of replacing conventional gas turbine APUs with SOFC/GT APUs on aircraft. An SOFC/GT design algorithm was created in order to determine the specifications of an SOFC/GT APU. The design algorithm is comprised of several integrated modules which together model the characteristics of each component of the SOFC/GT system. Given certain overall inputs, through numerical analysis, the algorithm produces an SOFC/GT APU, optimized for specific power and efficiency, capable of performing to the required specifications. The SOFC/GT design is then input into a previously developed quasi-dynamic SOFC/GT model to determine its load following capabilities over an aircraft flight cycle. Finally an aircraft range study is conducted to determine the feasibility of the SOFC/GT APU as a replacement for the conventional gas turbine APU. The design results show that SOFC/GT APUs have lower specific power than GT systems, but have much higher efficiencies. Moreover, the dynamic simulation results show that SOFC/GT APUs are capable of following modern flight loads. Finally, the range study determined that SOFC/GT APUs are more attractive over conventional APUs for longer range aircraft.
Aircraft energy efficiency laminar flow control wing design study
NASA Technical Reports Server (NTRS)
Bonner, T. F., Jr.; Pride, J. D., Jr.; Fernald, W. W.
1977-01-01
An engineering design study was performed in which laminar flow control (LFC) was integrated into the wing of a commercial passenger transport aircraft. A baseline aircraft configuration was selected and the wing geometry was defined. The LFC system, with suction slots, ducting, and suction pumps was integrated with the wing structure. The use of standard aluminum technology and advanced superplastic formed diffusion bonded titanium technology was evaluated. The results of the design study show that the LFC system can be integrated with the wing structure to provide a structurally and aerodynamically efficient wing for a commercial transport aircraft.
Aircraft conceptual design - an adaptable parametric sizing methodology
NASA Astrophysics Data System (ADS)
Coleman, Gary John, Jr.
Aerospace is a maturing industry with successful and refined baselines which work well for traditional baseline missions, markets and technologies. However, when new markets (space tourism) or new constrains (environmental) or new technologies (composite, natural laminar flow) emerge, the conventional solution is not necessarily best for the new situation. Which begs the question "how does a design team quickly screen and compare novel solutions to conventional solutions for new aerospace challenges?" The answer is rapid and flexible conceptual design Parametric Sizing. In the product design life-cycle, parametric sizing is the first step in screening the total vehicle in terms of mission, configuration and technology to quickly assess first order design and mission sensitivities. During this phase, various missions and technologies are assessed. During this phase, the designer is identifying design solutions of concepts and configurations to meet combinations of mission and technology. This research undertaking contributes the state-of-the-art in aircraft parametric sizing through (1) development of a dedicated conceptual design process and disciplinary methods library, (2) development of a novel and robust parametric sizing process based on 'best-practice' approaches found in the process and disciplinary methods library, and (3) application of the parametric sizing process to a variety of design missions (transonic, supersonic and hypersonic transports), different configurations (tail-aft, blended wing body, strut-braced wing, hypersonic blended bodies, etc.), and different technologies (composite, natural laminar flow, thrust vectored control, etc.), in order to demonstrate the robustness of the methodology and unearth first-order design sensitivities to current and future aerospace design problems. This research undertaking demonstrates the importance of this early design step in selecting the correct combination of mission, technologies and configuration to
OPTIMAL NETWORK TOPOLOGY DESIGN
NASA Technical Reports Server (NTRS)
Yuen, J. H.
1994-01-01
This program was developed as part of a research study on the topology design and performance analysis for the Space Station Information System (SSIS) network. It uses an efficient algorithm to generate candidate network designs (consisting of subsets of the set of all network components) in increasing order of their total costs, and checks each design to see if it forms an acceptable network. This technique gives the true cost-optimal network, and is particularly useful when the network has many constraints and not too many components. It is intended that this new design technique consider all important performance measures explicitly and take into account the constraints due to various technical feasibilities. In the current program, technical constraints are taken care of by the user properly forming the starting set of candidate components (e.g. nonfeasible links are not included). As subsets are generated, they are tested to see if they form an acceptable network by checking that all requirements are satisfied. Thus the first acceptable subset encountered gives the cost-optimal topology satisfying all given constraints. The user must sort the set of "feasible" link elements in increasing order of their costs. The program prompts the user for the following information for each link: 1) cost, 2) connectivity (number of stations connected by the link), and 3) the stations connected by that link. Unless instructed to stop, the program generates all possible acceptable networks in increasing order of their total costs. The program is written only to generate topologies that are simply connected. Tests on reliability, delay, and other performance measures are discussed in the documentation, but have not been incorporated into the program. This program is written in PASCAL for interactive execution and has been implemented on an IBM PC series computer operating under PC DOS. The disk contains source code only. This program was developed in 1985.
Advanced airfoil design empirically based transonic aircraft drag buildup technique
NASA Technical Reports Server (NTRS)
Morrison, W. D., Jr.
1976-01-01
To systematically investigate the potential of advanced airfoils in advance preliminary design studies, empirical relationships were derived, based on available wind tunnel test data, through which total drag is determined recognizing all major aircraft geometric variables. This technique recognizes a single design lift coefficient and Mach number for each aircraft. Using this technique drag polars are derived for all Mach numbers up to MDesign + 0.05 and lift coefficients -0.40 to +0.20 from CLDesign.
NASA Technical Reports Server (NTRS)
Skillen, Michael D.; Crossley, William A.
2008-01-01
This report presents an approach for sizing of a morphing aircraft based upon a multi-level design optimization approach. For this effort, a morphing wing is one whose planform can make significant shape changes in flight - increasing wing area by 50% or more from the lowest possible area, changing sweep 30 or more, and/or increasing aspect ratio by as much as 200% from the lowest possible value. The top-level optimization problem seeks to minimize the gross weight of the aircraft by determining a set of "baseline" variables - these are common aircraft sizing variables, along with a set of "morphing limit" variables - these describe the maximum shape change for a particular morphing strategy. The sub-level optimization problems represent each segment in the morphing aircraft's design mission; here, each sub-level optimizer minimizes fuel consumed during each mission segment by changing the wing planform within the bounds set by the baseline and morphing limit variables from the top-level problem.
NASA Technical Reports Server (NTRS)
Reuther, James; Alonso, Juan Jose; Rimlinger, Mark J.; Jameson, Antony
1996-01-01
This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. The design process is greatly accelerated through the use of both control theory and a parallel implementation on distributed memory computers. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods (13, 12, 44, 38). The resulting problem is then implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) Standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on higher order computational fluid dynamics methods (CFD). In our earlier studies, the serial implementation of this design method (19, 20, 21, 23, 39, 25, 40, 41, 42, 43, 9) was shown to be effective for the optimization of airfoils, wings, wing-bodies, and complex aircraft configurations using both the potential equation and the Euler equations (39, 25). In our most recent paper, the Euler method was extended to treat complete aircraft configurations via a new multiblock implementation. Furthermore, during the same conference, we also presented preliminary results demonstrating that the basic methodology could be ported to distributed memory parallel computing architectures [241. In this paper, our concem will be to demonstrate that the combined power of these new technologies can be used routinely in an industrial design environment by applying it to the case study of the design of typical supersonic transport configurations. A particular difficulty of this test case is posed by the propulsion/airframe integration.
NASA Technical Reports Server (NTRS)
Reuther, James; Alonso, Juan Jose; Rimlinger, Mark J.; Jameson, Antony
1996-01-01
This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. The design process is greatly accelerated through the use of both control theory and a parallel implementation on distributed memory computers. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods. The resulting problem is then implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) Standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on higher order computational fluid dynamics methods (CFD). In our earlier studies, the serial implementation of this design method was shown to be effective for the optimization of airfoils, wings, wing-bodies, and complex aircraft configurations using both the potential equation and the Euler equations. In our most recent paper, the Euler method was extended to treat complete aircraft configurations via a new multiblock implementation. Furthermore, during the same conference, we also presented preliminary results demonstrating that this basic methodology could be ported to distributed memory parallel computing architectures. In this paper, our concern will be to demonstrate that the combined power of these new technologies can be used routinely in an industrial design environment by applying it to the case study of the design of typical supersonic transport configurations. A particular difficulty of this test case is posed by the propulsion/airframe integration.
Integrated structural-aerodynamic design optimization
NASA Technical Reports Server (NTRS)
Haftka, R. T.; Kao, P. J.; Grossman, B.; Polen, D.; Sobieszczanski-Sobieski, J.
1988-01-01
This paper focuses on the processes of simultaneous aerodynamic and structural wing design as a prototype for design integration, with emphasis on the major difficulty associated with multidisciplinary design optimization processes, their enormous computational costs. Methods are presented for reducing this computational burden through the development of efficient methods for cross-sensitivity calculations and the implementation of approximate optimization procedures. Utilizing a modular sensitivity analysis approach, it is shown that the sensitivities can be computed without the expensive calculation of the derivatives of the aerodynamic influence coefficient matrix, and the derivatives of the structural flexibility matrix. The same process is used to efficiently evaluate the sensitivities of the wing divergence constraint, which should be particularly useful, not only in problems of complete integrated aircraft design, but also in aeroelastic tailoring applications.
Conceptual design for a laminar-flying-wing aircraft
NASA Astrophysics Data System (ADS)
Saeed, T. I.
The laminar-flying-wing aircraft appears to be an attractive long-term prospect for reducing the environmental impact of commercial aviation. In assessing its potential, a relatively straightforward initial step is the conceptual design of a version with restricted sweep angle. Such a design is the topic of this thesis. Subject to constraints, this research aims to; provide insight into the parameters affecting practical laminar-flow-control suction power requirements; identify a viable basic design specification; and, on the basis of this, an assessment of the fuel efficiency through a detailed conceptual design study. It is shown that there is a minimum power requirement independent of the suction system design, associated with the stagnation pressure loss in the boundary layer. This requirement increases with aerofoil section thickness, but depends only weakly on Mach number and (for a thick, lightly-loaded laminar flying wing) lift coefficient. Deviation from the optimal suction distribution, due to a practical chamber-based architecture, is found to have very little effect on the overall suction coefficient. In the spanwise direction, through suitable choice of chamber depth, the pressure drop due to frictional and inertial effects may be rendered negligible. Finally, it is found that the pressure drop from the aerofoil surface to the pump collector ducts determines the power penalty. To identify the viable basic design specification, a high-level exploration of the laminar flying wing design space is performed. The characteristics of the design are assessed as a function of three parameters: thickness-to-chord ratio, wingspan, and unit Reynolds number. A feasible specification, with 20% thickness-to-chord, 80 m span and a unit Reynolds number of 8 x 106 m-1, is identified; it corresponds to a 187 tonne aircraft which cruises at Mach 0.67 and altitude 22,500 ft, with lift coefficient 0.14. On the basis of this specification, a detailed conceptual design is
Preliminary design studies of an advanced general aviation aircraft
NASA Technical Reports Server (NTRS)
Barrett, Ron; Demoss, Shane; Dirkzwager, AB; Evans, Darryl; Gomer, Charles; Keiter, Jerry; Knipp, Darren; Seier, Glen; Smith, Steve; Wenninger, ED
1991-01-01
The preliminary design results are presented of the advanced aircraft design project. The goal was to take a revolutionary look into the design of a general aviation aircraft. Phase 1 of the project included the preliminary design of two configurations, a pusher, and a tractor. Phase 2 included the selection of only one configuration for further study. The pusher configuration was selected on the basis of performance characteristics, cabin noise, natural laminar flow, and system layouts. The design was then iterated to achieve higher levels of performance.
Conceptual design of a flying boom for air-to-air refueling of passenger aircraft
NASA Astrophysics Data System (ADS)
Timmermans, Ir. H. S.; La Rocca, ir. G., Dr.
2014-10-01
This paper describes the conceptual development of a flying boom for air-to-air refuelingof passenger aircraft. This operational concept is currently evaluated within the EC project RECREATE as a possible means to achieve significant increase in overall fuel efficiency. While in military aviation aerial refueling is performed with the tankerflyingahead and above the receiver aircraft, in case of passenger aircraft, safety, cost and comfort criteria suggest to invert the set up. This unconventional configuration would require a different refueling boom, able to extend from the tanker towards the cruiser, against wind and gravity. Amultidisciplinary design optimization framework was set up to size and compare various boom design solutions free of structural divergence and sufficientlycontrollable and with minimum values of weight and drag. Oneconcept, based on an innovative kinematic mechanism, was selected for its ability to meet all design constraints, with weight and drag values comparable to conventional boom designs.
Profile design for an advanced-technology airfoil for general aviation aircraft
NASA Technical Reports Server (NTRS)
Welte, D.
1978-01-01
A profile from the NASA General Aviation Whitcomb series and NACA profiles are used as a starting point in designing an advanced airfoil for general aviation aircraft. Potential theory pressure distribution calculations, together with boundary layer calculations, permit a decrease in the null moment and an optimization of the lift characteristics of the wing. Trailing edge flap design is also improved. Wind tunnel tests are used to compare the conventional profiles, the NASA profile, and the improved design.
V/STOL tilt rotor aircraft study. Volume 2: Preliminary design of research aircraft
NASA Technical Reports Server (NTRS)
1972-01-01
A preliminary design study was conducted to establish a minimum sized, low cost V/STOL tilt-rotor research aircraft with the capability of performing proof-of-concept flight research investigations applicable to a wide range of useful military and commercial configurations. The analysis and design approach was based on state-of-the-art methods and maximum use of off-the-shelf hardware and systems to reduce development risk, procurement cost and schedules impact. The rotors to be used are of 26 foot diameter and are the same as currently under construction and test as part of NASA Tilt-Rotor Contract NAS2-6505. The aircraft has a design gross weight of 12,000 lbs. The proposed engines to be used are Lycoming T53-L-13B rated at 1550 shaft horsepower which are fully qualified. A flight test investigation is recommended which will determine the capabilities and limitations of the research aircraft.
An analytical sensitivity method for use in integrated aeroservoelastic aircraft design
NASA Technical Reports Server (NTRS)
Gilbert, Michael G.
1989-01-01
Interdisciplinary analysis capabilities have been developed for aeroservoelastic aircraft and large flexible spacecraft, but the requisite integrated design methods are only beginning to be developed. One integrated design method which has received attention is based on hierarchal problem decompositions, optimization, and design sensitivity analyses. This paper highlights a design sensitivity analysis method for Linear Quadratic Gaussian (LQG) optimal control laws, enabling the use of LQG techniques in the hierarchal design methodology. The LQG sensitivity analysis method calculates the change in the optimal control law and resulting controlled system responses due to changes in fixed design integration parameters using analytical sensitivity equations. Numerical results of an LQG design sensitivity analysis for a realistic aeroservoelastic aircraft example are presented. In this example, the sensitivity of the optimal control law and aircraft response for various parameters such as wing bending natural frequency is determined. The sensitivity results computed from the analytical expressions are used to estimate changes in response resulting from changes in the parameters. Comparisons of the estimates with exact calculated responses show they are reasonably accurate for + or - 15 percent changes in the parameters. Evaluation of the analytical expressions is computationally faster than equivalent finite difference calculations.
An analytical sensitivity method for use in integrated aeroservoelastic aircraft design
NASA Technical Reports Server (NTRS)
Gilbert, Michael G.
1989-01-01
Interdisciplinary analysis capabilities have been developed for aeroservoelastic aircraft and large flexible spacecraft, but the requisite integrated design methods are only beginning to be developed. One integrated design method which has received attention is based on hierarchal problem decompositions, optimization, and design sensitivity analyses. This paper highlights a design sensitivity analysis method for Linear Quadratic Gaussian (LQG) optimal control laws, enabling the use of LQG techniques in the hierarchal design methodology. The LQG sensitivity analysis method calculates the change in the optimal control law and resulting controlled system responses due to changes in fixed design integration parameters using analytical sensitivity equations. Numerical results of a LQG design sensitivity analysis for a realistic aeroservoelastic aircraft example are presented. In this example, the sensitivity of the optimal control law and aircraft response for various parameters such as wing bending natural frequency is determined. The sensitivity results computed from the analytical expressions are used to estimate changes in response resulting from changes in the parameters. Comparisons of the estimates with exact calculated responses show they are reasonably accurate for + or - 15 percent changes in the parameters. Evaluation of the analytical expressions is computationally faster than equivalent finite difference calculations.
Design synthesis and optimization of joined-wing transports
NASA Technical Reports Server (NTRS)
Gallman, John W.; Smith, Stephen C.; Kroo, Ilan M.
1990-01-01
A computer program for aircraft synthesis using a numerical optimizer was developed to study the application of the joined-wing configuration to transport aircraft. The structural design algorithm included the effects of secondary bending moments to investigate the possibility of tail buckling and to design joined wings resistant to buckling. The structural weight computed using this method was combined with a statistically-based method to obtain realistic estimates of total lifting surface weight and aircraft empty weight. A variety of 'optimum' joined-wing and conventional aircraft designs were compared on the basis of direct operating cost, gross weight, and cruise drag. The most promising joined-wing designs were found to have a joint location at about 70 percent of the wing semispan. The optimum joined-wing transport is shown to save 1.7 percent in direct operating cost and 11 percent in drag for a 2000 nautical mile transport mission.
Simple uncertainty propagation for early design phase aircraft sizing
NASA Astrophysics Data System (ADS)
Lenz, Annelise
Many designers and systems analysts are aware of the uncertainty inherent in their aircraft sizing studies; however, few incorporate methods to address and quantify this uncertainty. Many aircraft design studies use semi-empirical predictors based on a historical database and contain uncertainty -- a portion of which can be measured and quantified. In cases where historical information is not available, surrogate models built from higher-fidelity analyses often provide predictors for design studies where the computational cost of directly using the high-fidelity analyses is prohibitive. These surrogate models contain uncertainty, some of which is quantifiable. However, rather than quantifying this uncertainty, many designers merely include a safety factor or design margin in the constraints to account for the variability between the predicted and actual results. This can become problematic if a designer does not estimate the amount of variability correctly, which then can result in either an "over-designed" or "under-designed" aircraft. "Under-designed" and some "over-designed" aircraft will likely require design changes late in the process and will ultimately require more time and money to create; other "over-designed" aircraft concepts may not require design changes, but could end up being more costly than necessary. Including and propagating uncertainty early in the design phase so designers can quantify some of the errors in the predictors could help mitigate the extent of this additional cost. The method proposed here seeks to provide a systematic approach for characterizing a portion of the uncertainties that designers are aware of and propagating it throughout the design process in a procedure that is easy to understand and implement. Using Monte Carlo simulations that sample from quantified distributions will allow a systems analyst to use a carpet plot-like approach to make statements like: "The aircraft is 'P'% likely to weigh 'X' lbs or less, given the
An economic model for evaluating high-speed aircraft designs
NASA Technical Reports Server (NTRS)
Vandervelden, Alexander J. M.
1989-01-01
A Class 1 method for determining whether further development of a new aircraft design is desirable from all viewpoints is presented. For the manufacturer the model gives an estimate of the total cost of research and development from the preliminary design to the first production aircraft. Using Wright's law of production, one can derive the average cost per aircraft produced for a given break-even number. The model will also provide the airline with a good estimate of the direct and indirect operating costs. From the viewpoint of the passenger, the model proposes a tradeoff between ticket price and cruise speed. Finally all of these viewpoints are combined in a Comparative Aircraft Seat-kilometer Economic Index.
Needs and Challenges in Education for Aircraft Design.
ERIC Educational Resources Information Center
Haupt, Ulrich
A brief review of recent developments in engineering education leads to basic reflections about the importance of design education. Aircraft design is singled out as a field where demands on design are particularly high and urgent. Basic needs are determined. Additional challenges posed by engineering technology, continuing studies,â€¦
Aerodynamic design using numerical optimization
NASA Technical Reports Server (NTRS)
Murman, E. M.; Chapman, G. T.
1983-01-01
The procedure of using numerical optimization methods coupled with computational fluid dynamic (CFD) codes for the development of an aerodynamic design is examined. Several approaches that replace wind tunnel tests, develop pressure distributions and derive designs, or fulfill preset design criteria are presented. The method of Aerodynamic Design by Numerical Optimization (ADNO) is described and illustrated with examples.
Aircraft integrated design and analysis: A classroom experience
NASA Technical Reports Server (NTRS)
Weisshaar, Terrence A.
1989-01-01
AAE 451 is the capstone course required of all senior undergraduates in the School of Aeronautics and Astronautics at Purdue University. During the past year the first steps of a long evolutionary process were taken to change the content and expectations of this course. These changes are the result of the availability of advanced computational capabilities and sophisticated electronic media availability at Purdue. This presentation will describe both the long range objectives and this year's experience using the High Speed Commercial Transport design, the AIAA Long Duration Aircraft design and RPV design proposal as project objectives. The central goal of these efforts is to provide a user-friendly, computer-software-based environment to supplement traditional design course methodology. The Purdue University Computer Center (PUCC), the Engineering Computer Network (ECN) and stand-alone PC's are being used for this development. This year's accomplishments center primarily on aerodynamics software obtained from NASA/Langley and its integration into the classroom. Word processor capability for oral and written work and computer graphics were also blended into the course. A total of ten HSCT designs were generated, ranging from twin-fuselage aircraft, forward swept wing aircraft to the more traditional delta and double-delta wing aircraft. Four Long Duration Aircraft designs were submitted, together with one RPV design tailored for photographic surveillance.
Structural Optimization in automotive design
NASA Technical Reports Server (NTRS)
Bennett, J. A.; Botkin, M. E.
1984-01-01
Although mathematical structural optimization has been an active research area for twenty years, there has been relatively little penetration into the design process. Experience indicates that often this is due to the traditional layout-analysis design process. In many cases, optimization efforts have been outgrowths of analysis groups which are themselves appendages to the traditional design process. As a result, optimization is often introduced into the design process too late to have a significant effect because many potential design variables have already been fixed. A series of examples are given to indicate how structural optimization has been effectively integrated into the design process.
Aircraft systems design studies employing advanced transport technologies
NASA Technical Reports Server (NTRS)
Downie, B.; Pearce, C.; Quartero, C.; Taylor, A.
1972-01-01
System and design integration studies are presented to define and assess the application of the advanced technology most likely to result in a superior next generation, high subsonic/sonic conventional takeoff and landing transport aircraft system. It is concluded that the new technologies can be directed toward the achievement of improved economy and performance. These benefits may be used to compensate for the penalties associated with reduced noise requirements anticipated to make future aircraft ecologically acceptable.
Control Design for a Generic Commercial Aircraft Engine
NASA Technical Reports Server (NTRS)
Csank, Jeffrey; May, Ryan D.
2010-01-01
This paper describes the control algorithms and control design process for a generic commercial aircraft engine simulation of a 40,000 lb thrust class, two spool, high bypass ratio turbofan engine. The aircraft engine is a complex nonlinear system designed to operate over an extreme range of environmental conditions, at temperatures from approximately -60 to 120+ F, and at altitudes from below sea level to 40,000 ft, posing multiple control design constraints. The objective of this paper is to provide the reader an overview of the control design process, design considerations, and justifications as to why the particular architecture and limits have been chosen. The controller architecture contains a gain-scheduled Proportional Integral controller along with logic to protect the aircraft engine from exceeding any limits. Simulation results illustrate that the closed loop system meets the Federal Aviation Administration s thrust response requirements
QCGAT aircraft/engine design for reduced noise and emissions
NASA Technical Reports Server (NTRS)
Lanson, L.; Terrill, K. M.
1980-01-01
The high bypass ratio QCGAT engine played an important role in shaping the aircraft design. The aircraft which evolved is a sleek, advanced design, six-place aircraft with 3538 kg (7,800 lb) maximum gross weight. It offers a 2778 kilometer (1500 nautical mile) range with cruise speed of 0.5 Mach number and will take-off and land on the vast majority of general aviation airfields. Advanced features include broad application of composite materials and a supercritical wing design with winglets. Full-span fowler flaps were introduced to improve landing capability. Engines are fuselage-mounted with inlets over the wing to provide shielding of fan noise by the wing surfaces. The design objectives, noise, and emission considerations, engine cycle and engine description are discussed as well as specific design features.
NASA Technical Reports Server (NTRS)
Chan, Agnes; Conley, Kristin; Javorski, Christian T.; Cheung, Kwok-Hung; Crivelli, Paul M.; Torrey, Nancy P.; Traver, Michael L.
1992-01-01
Increase in energy demands coupled with rapid depletion of natural energy resources have deemed solar energy as the most logical alternative source of power. The major objective of this project was to build a solar powered remotely controlled aircraft to demonstrate the feasibility of solar energy as an effective, alternate source of power. The final design was optimized for minimum weight and maximum strength of the structure. These design constraints necessitated a carbon fiber composite structure. Surya is a lightweight, durable aircraft capable of achieving level flight powered entirely by solar cells.
Lessons Learned During Solutions of Multidisciplinary Design Optimization Problems
NASA Technical Reports Server (NTRS)
Patnaik, Suna N.; Coroneos, Rula M.; Hopkins, Dale A.; Lavelle, Thomas M.
2000-01-01
Optimization research at NASA Glenn Research Center has addressed the design of structures, aircraft and airbreathing propulsion engines. During solution of the multidisciplinary problems several issues were encountered. This paper lists four issues and discusses the strategies adapted for their resolution: (1) The optimization process can lead to an inefficient local solution. This deficiency was encountered during design of an engine component. The limitation was overcome through an augmentation of animation into optimization. (2) Optimum solutions obtained were infeasible for aircraft and air-breathing propulsion engine problems. Alleviation of this deficiency required a cascading of multiple algorithms. (3) Profile optimization of a beam produced an irregular shape. Engineering intuition restored the regular shape for the beam. (4) The solution obtained for a cylindrical shell by a subproblem strategy converged to a design that can be difficult to manufacture. Resolution of this issue remains a challenge. The issues and resolutions are illustrated through six problems: (1) design of an engine component, (2) synthesis of a subsonic aircraft, (3) operation optimization of a supersonic engine, (4) design of a wave-rotor-topping device, (5) profile optimization of a cantilever beam, and (6) design of a cvlindrical shell. The combined effort of designers and researchers can bring the optimization method from academia to industry.
The ARCTAS aircraft mission: design and execution
NASA Astrophysics Data System (ADS)
Jacob, D. J.; Crawford, J. H.; Maring, H. B.; Clarke, A. D.; Dibb, J. E.; Ferrare, R. A.; Hostetler, C. A.; Russell, P. B.; Singh, H. B.; Thompson, A. M.; Shaw, G. E.; McCauley, E.; Pederson, J. R.; Fisher, J. A.
2009-12-01
We present an overview of the NASA Arctic Research of the Composition of the Troposphere from Aircraft and Satellites (ARCTAS) mission, conducted in two 3-week deployments based in Alaska (April 2008) and western Canada (June-July 2008). The goal of ARCTAS was to better understand the factors driving current changes in Arctic atmospheric composition and climate, including (1) transport of mid-latitude pollution, (2) boreal forest fires, (3) aerosol radiative forcing, and (4) chemical processes. ARCTAS involved three aircraft: a DC-8 with detailed chemical payload, a P-3 with extensive aerosol payload, and a B-200 with aerosol remote sensing instrumentation. The aircraft augmented satellite observations of Arctic atmospheric composition, in particular from the NASA A-Train, by (1) validating the data, (2) improving constraints on retrievals, (3) making correlated observations, and (4) characterizing chemical and aerosol processes. The April flights (ARCTAS-A) sampled pollution plumes from all three mid-latitude continents, fire plumes from Siberia and Southeast Asia, and halogen radical events. The June-July flights (ARCTAS-B) focused on boreal forest fire influences and sampled fresh fire plumes from northern Saskatchewan as well as older fire plumes from Canada, Siberia, and California. The June-July deployment was preceded by one week of flights over California sponsored by the California Air Resources Board (ARCTAS-CARB). The ARCTAS-CARB goals were to (1) improve state emission inventories for greenhouse gases and aerosols, (2) provide observations to test and improve models of ozone and aerosol pollution. Extensive sampling across southern California and the Central Valley characterized emissions from urban centers, offshore shipping lanes, agricultural crops, feedlots, industrial sources, and wildfires.
Preliminary design studies of an advanced general aviation aircraft
NASA Technical Reports Server (NTRS)
1991-01-01
Preliminary design studies are presented for an advanced general aviation aircraft. Advanced guidance and display concepts, laminar flow, smart structures, fuselage and wing structural design and manufacturing, and preliminary configuration design are discussed. This project was conducted as a graduate level design class under the auspices of the KU/NASA/USRA Advanced Design Program in Aeronautics. The results obtained during the fall semester of 1990 (Phase 1) and the spring semester of 1991 (Phase 2) are presented.
Integrated controls design optimization
Lou, Xinsheng; Neuschaefer, Carl H.
2015-09-01
A control system (207) for optimizing a chemical looping process of a power plant includes an optimizer (420), an income algorithm (230) and a cost algorithm (225) and a chemical looping process models. The process models are used to predict the process outputs from process input variables. Some of the process in puts and output variables are related to the income of the plant; and some others are related to the cost of the plant operations. The income algorithm (230) provides an income input to the optimizer (420) based on a plurality of input parameters (215) of the power plant. The cost algorithm (225) provides a cost input to the optimizer (420) based on a plurality of output parameters (220) of the power plant. The optimizer (420) determines an optimized operating parameter solution based on at least one of the income input and the cost input, and supplies the optimized operating parameter solution to the power plant.
Advances in Experiment Design for High Performance Aircraft
NASA Technical Reports Server (NTRS)
Morelli, Engene A.
1998-01-01
A general overview and summary of recent advances in experiment design for high performance aircraft is presented, along with results from flight tests. General theoretical background is included, with some discussion of various approaches to maneuver design. Flight test examples from the F-18 High Alpha Research Vehicle (HARV) are used to illustrate applications of the theory. Input forms are compared using Cramer-Rao bounds for the standard errors of estimated model parameters. Directions for future research in experiment design for high performance aircraft are identified.
Hansborough, L.; Hamm, R.; Stovall, J.; Swenson, D.
1980-01-01
PIGMI (Pion Generator for Medical Irradiations) is a compact linear proton accelerator design, optimized for pion production and cancer treatment use in a hospital environment. Technology developed during a four-year PIGMI Prototype experimental program allows the design of smaller, less expensive, and more reliable proton linacs. A new type of low-energy accelerating structure, the radio-frequency quadrupole (RFQ) has been tested; it produces an exceptionally good-quality beam and allows the use of a simple 30-kV injector. Average axial electric-field gradients of over 9 MV/m have been demonstrated in a drift-tube linac (DTL) structure. Experimental work is underway to test the disk-and-washer (DAW) structure, another new type of accelerating structure for use in the high-energy coupled-cavity linac (CCL). Sufficient experimental and developmental progress has been made to closely define an actual PIGMI. It will consist of a 30-kV injector, and RFQ linac to a proton energy of 2.5 MeV, a DTL linac to 125 MeV, and a CCL linac to the final energy of 650 MeV. The total length of the accelerator is 133 meters. The RFQ and DTL will be driven by a single 440-MHz klystron; the CCL will be driven by six 1320-MHz klystrons. The peak beam current is 28 mA. The beam pulse length is 60 ..mu..s at a 60-Hz repetition rate, resulting in a 100-..mu..A average beam current. The total cost of the accelerator is estimated to be approx. $10 million.
Multidisciplinary design optimization using genetic algorithms
NASA Astrophysics Data System (ADS)
Unal, Resit
1994-12-01
Multidisciplinary design optimization (MDO) is an important step in the conceptual design and evaluation of launch vehicles since it can have a significant impact on performance and life cycle cost. The objective is to search the system design space to determine values of design variables that optimize the performance characteristic subject to system constraints. Gradient-based optimization routines have been used extensively for aerospace design optimization. However, one limitation of gradient based optimizers is their need for gradient information. Therefore, design problems which include discrete variables can not be studied. Such problems are common in launch vehicle design. For example, the number of engines and material choices must be integer values or assume only a few discrete values. In this study, genetic algorithms are investigated as an approach to MDO problems involving discrete variables and discontinuous domains. Optimization by genetic algorithms (GA) uses a search procedure which is fundamentally different from those gradient based methods. Genetic algorithms seek to find good solutions in an efficient and timely manner rather than finding the best solution. GA are designed to mimic evolutionary selection. A population of candidate designs is evaluated at each iteration, and each individual's probability of reproduction (existence in the next generation) depends on its fitness value (related to the value of the objective function). Progress toward the optimum is achieved by the crossover and mutation operations. GA is attractive since it uses only objective function values in the search process, so gradient calculations are avoided. Hence, GA are able to deal with discrete variables. Studies report success in the use of GA for aircraft design optimization studies, trajectory analysis, space structure design and control systems design. In these studies reliable convergence was achieved, but the number of function evaluations was large compared
Multidisciplinary design optimization using genetic algorithms
NASA Technical Reports Server (NTRS)
Unal, Resit
1994-01-01
Multidisciplinary design optimization (MDO) is an important step in the conceptual design and evaluation of launch vehicles since it can have a significant impact on performance and life cycle cost. The objective is to search the system design space to determine values of design variables that optimize the performance characteristic subject to system constraints. Gradient-based optimization routines have been used extensively for aerospace design optimization. However, one limitation of gradient based optimizers is their need for gradient information. Therefore, design problems which include discrete variables can not be studied. Such problems are common in launch vehicle design. For example, the number of engines and material choices must be integer values or assume only a few discrete values. In this study, genetic algorithms are investigated as an approach to MDO problems involving discrete variables and discontinuous domains. Optimization by genetic algorithms (GA) uses a search procedure which is fundamentally different from those gradient based methods. Genetic algorithms seek to find good solutions in an efficient and timely manner rather than finding the best solution. GA are designed to mimic evolutionary selection. A population of candidate designs is evaluated at each iteration, and each individual's probability of reproduction (existence in the next generation) depends on its fitness value (related to the value of the objective function). Progress toward the optimum is achieved by the crossover and mutation operations. GA is attractive since it uses only objective function values in the search process, so gradient calculations are avoided. Hence, GA are able to deal with discrete variables. Studies report success in the use of GA for aircraft design optimization studies, trajectory analysis, space structure design and control systems design. In these studies reliable convergence was achieved, but the number of function evaluations was large compared
Design Optimization Toolkit: Users' Manual
Aguilo Valentin, Miguel Alejandro
2014-07-01
The Design Optimization Toolkit (DOTk) is a stand-alone C++ software package intended to solve complex design optimization problems. DOTk software package provides a range of solution methods that are suited for gradient/nongradient-based optimization, large scale constrained optimization, and topology optimization. DOTk was design to have a flexible user interface to allow easy access to DOTk solution methods from external engineering software packages. This inherent flexibility makes DOTk barely intrusive to other engineering software packages. As part of this inherent flexibility, DOTk software package provides an easy-to-use MATLAB interface that enables users to call DOTk solution methods directly from the MATLAB command window.
Automatic carrier landing system for V/STOL aircraft using L1 adaptive and optimal control
NASA Astrophysics Data System (ADS)
Hariharapura Ramesh, Shashank
This thesis presents a framework for developing automatic carrier landing systems for aircraft with vertical or short take-off and landing capability using two different control strategies---gain-scheduled linear optimal control, and L1 adaptive control. The carrier landing sequence of V/STOL aircraft involves large variations in dynamic pressure and aerodynamic coefficients arising because of the transition from aerodynamic-supported to jet-borne flight, descent to the touchdown altitude, and turns performed to align with the runway. Consequently, the dynamics of the aircraft exhibit a highly non-linear dynamical behavior with variations in flight conditions prior to touchdown. Therefore, the implication is the need for non-linear control techniques to achieve automatic landing. Gain-scheduling has been one of the most widely employed techniques for control of aircraft, which involves designing linear controllers for numerous trimmed flight conditions, and interpolating them to achieve a global non-linear control. Adaptive control technique, on the other hand, eliminates the need to schedule the controller parameters as they adapt to changing flight conditions.
Engine Yaw Augmentation for Hybrid-Wing-Body Aircraft via Optimal Control Allocation Techniques
NASA Technical Reports Server (NTRS)
Taylor, Brian R.; Yoo, Seung-Yeun
2011-01-01
Asymmetric engine thrust was implemented in a hybrid-wing-body non-linear simulation to reduce the amount of aerodynamic surface deflection required for yaw stability and control. Hybrid-wing-body aircraft are especially susceptible to yaw surface deflection due to their decreased bare airframe yaw stability resulting from the lack of a large vertical tail aft of the center of gravity. Reduced surface deflection, especially for trim during cruise flight, could reduce the fuel consumption of future aircraft. Designed as an add-on, optimal control allocation techniques were used to create a control law that tracks total thrust and yaw moment commands with an emphasis on not degrading the baseline system. Implementation of engine yaw augmentation is shown and feasibility is demonstrated in simulation with a potential drag reduction of 2 to 4 percent. Future flight tests are planned to demonstrate feasibility in a flight environment.
Application of Adjoint Methodology to Supersonic Aircraft Design Using Reversed Equivalent Areas
NASA Technical Reports Server (NTRS)
Rallabhandi, Sriram K.
2013-01-01
This paper presents an approach to shape an aircraft to equivalent area based objectives using the discrete adjoint approach. Equivalent areas can be obtained either using reversed augmented Burgers equation or direct conversion of off-body pressures into equivalent area. Formal coupling with CFD allows computation of sensitivities of equivalent area objectives with respect to aircraft shape parameters. The exactness of the adjoint sensitivities is verified against derivatives obtained using the complex step approach. This methodology has the benefit of using designer-friendly equivalent areas in the shape design of low-boom aircraft. Shape optimization results with equivalent area cost functionals are discussed and further refined using ground loudness based objectives.
Practical input optimization for aircraft parameter estimation experiments. Ph.D. Thesis, 1990
NASA Technical Reports Server (NTRS)
Morelli, Eugene A.
1993-01-01
The object of this research was to develop an algorithm for the design of practical, optimal flight test inputs for aircraft parameter estimation experiments. A general, single pass technique was developed which allows global optimization of the flight test input design for parameter estimation using the principles of dynamic programming with the input forms limited to square waves only. Provision was made for practical constraints on the input, including amplitude constraints, control system dynamics, and selected input frequency range exclusions. In addition, the input design was accomplished while imposing output amplitude constraints required by model validity and considerations of safety during the flight test. The algorithm has multiple input design capability, with optional inclusion of a constraint that only one control move at a time, so that a human pilot can implement the inputs. It is shown that the technique can be used to design experiments for estimation of open loop model parameters from closed loop flight test data. The report includes a new formulation of the optimal input design problem, a description of a new approach to the solution, and a summary of the characteristics of the algorithm, followed by three example applications of the new technique which demonstrate the quality and expanded capabilities of the input designs produced by the new technique. In all cases, the new input design approach showed significant improvement over previous input design methods in terms of achievable parameter accuracies.
Eagle RTS: A design for a regional transport aircraft
NASA Technical Reports Server (NTRS)
Bryer, Paul; Buckles, Jon; Lemke, Paul; Peake, Kirk
1992-01-01
This university design project concerns the Eagle RTS (Regional Transport System), a 66 passenger, twin turboprop aircraft with a range of 836 nautical miles. It will operate with a crew of two pilots and two flight attendents. This aircraft will employ the use of aluminum alloys and composite materials to reduce the aircraft weight and increase aerodynamic efficiency. The Eagle RTS will use narrow body aerodynamics with a canard configuration to improve performance. Leading edge technology will be used in the cockpit to improve flight handling and safety. The Eagle RTS propulsion system will consist of two turboprop engines with a total thrust of approximately 6300 pounds, 3150 pounds thrust per engine, for the cruise configuration. The engines will be mounted on the aft section of the aircraft to increase passenger safety in the event of a propeller failure. Aft mounted engines will also increase the overall efficiency of the aircraft by reducing the aircraft's drag. The Eagle RTS is projected to have a takeoff distance of approximately 4700 feet and a landing distance of 6100 feet. These distances will allow the Eagle RTS to land at the relatively short runways of regional airports.
The Computer Aided Aircraft-design Package (CAAP)
NASA Technical Reports Server (NTRS)
Yalif, Guy U.
1994-01-01
The preliminary design of an aircraft is a complex, labor-intensive, and creative process. Since the 1970's, many computer programs have been written to help automate preliminary airplane design. Time and resource analyses have identified, 'a substantial decrease in project duration with the introduction of an automated design capability'. Proof-of-concept studies have been completed which establish 'a foundation for a computer-based airframe design capability', Unfortunately, today's design codes exist in many different languages on many, often expensive, hardware platforms. Through the use of a module-based system architecture, the Computer aided Aircraft-design Package (CAAP) will eventually bring together many of the most useful features of existing programs. Through the use of an expert system, it will add an additional feature that could be described as indispensable to entry level engineers and students: the incorporation of 'expert' knowledge into the automated design process.
A concept for adaptive performance optimization on commercial transport aircraft
NASA Technical Reports Server (NTRS)
Jackson, Michael R.; Enns, Dale F.
1995-01-01
An adaptive control method is presented for the minimization of drag during flight for transport aircraft. The minimization of drag is achieved by taking advantage of the redundant control capability available in the pitch axis, with the horizontal tail used as the primary surface and symmetric deflection of the ailerons and cruise flaps used as additional controls. The additional control surfaces are excited with sinusoidal signals, while the altitude and velocity loops are closed with guidance and control laws. A model of the throttle response as a function of the additional control surfaces is formulated and the parameters in the model are estimated from the sensor measurements using a least squares estimation method. The estimated model is used to determine the minimum drag positions of the control surfaces. The method is presented for the optimization of one and two additional control surfaces. The adaptive control method is extended to optimize rate of climb with the throttle fixed. Simulations that include realistic disturbances are presented, as well as the results of a Monte Carlo simulation analysis that shows the effects of changing the disturbance environment and the excitation signal parameters.
Aerodynamic Shape Optimization of Complex Aircraft Configurations via an Adjoint Formulation
NASA Technical Reports Server (NTRS)
Reuther, James; Jameson, Antony; Farmer, James; Martinelli, Luigi; Saunders, David
1996-01-01
This work describes the implementation of optimization techniques based on control theory for complex aircraft configurations. Here control theory is employed to derive the adjoint differential equations, the solution of which allows for a drastic reduction in computational costs over previous design methods (13, 12, 43, 38). In our earlier studies (19, 20, 22, 23, 39, 25, 40, 41, 42) it was shown that this method could be used to devise effective optimization procedures for airfoils, wings and wing-bodies subject to either analytic or arbitrary meshes. Design formulations for both potential flows and flows governed by the Euler equations have been demonstrated, showing that such methods can be devised for various governing equations (39, 25). In our most recent works (40, 42) the method was extended to treat wing-body configurations with a large number of mesh points, verifying that significant computational savings can be gained for practical design problems. In this paper the method is extended for the Euler equations to treat complete aircraft configurations via a new multiblock implementation. New elements include a multiblock-multigrid flow solver, a multiblock-multigrid adjoint solver, and a multiblock mesh perturbation scheme. Two design examples are presented in which the new method is used for the wing redesign of a transonic business jet.
The Guardian: Preliminary design of a close air support aircraft
NASA Technical Reports Server (NTRS)
Haag, Jonathan; Huber, David; Mcinerney, Kelly; Mulligan, Greg; Pessin, David; Seelos, Michael
1991-01-01
One design is presented of a Close Air Support (CAS) aircraft. It is a canard wing, twin engine, twin vertical tail aircraft that has the capability to cruise at 520 knots. The Guardian contains state of the art flight control systems. Specific highlights of the Guardian include: (1) low cost (the acquisition cost per airplane is $13.6 million for a production of 500 airplanes); (2) low maintenance (it was designed to be easily maintainable in unprepared fields); and (3) high versatility (it can perform a wide range of missions). Along with being a CAS aircraft, it is capable of long ferry missions, battlefield interdiction, maritime attack, and combat rescue. The Guardian is capable of a maximum ferry of 3800 nm, can takeoff in a distance of 1700 ft, land in a ground roll distance of 1644 ft. It has a maximum takeoff weight of 48,753 lbs, and is capable of carrying up to 19,500 lbs of ordinance.
Preliminary aerodynamic design considerations for advanced laminar flow aircraft configurations
NASA Technical Reports Server (NTRS)
Johnson, Joseph L., Jr.; Yip, Long P.; Jordan, Frank L., Jr.
1986-01-01
Modern composite manufacturing methods have provided the opportunity for smooth surfaces that can sustain large regions of natural laminar flow (NLF) boundary layer behavior and have stimulated interest in developing advanced NLF airfoils and improved aircraft designs. Some of the preliminary results obtained in exploratory research investigations on advanced aircraft configurations at the NASA Langley Research Center are discussed. Results of the initial studies have shown that the aerodynamic effects of configuration variables such as canard/wing arrangements, airfoils, and pusher-type and tractor-type propeller installations can be particularly significant at high angles of attack. Flow field interactions between aircraft components were shown to produce undesirable aerodynamic effects on a wing behind a heavily loaded canard, and the use of properly designed wing leading-edge modifications, such as a leading-edge droop, offset the undesirable aerodynamic effects by delaying wing stall and providing increased stall/spin resistance with minimum degradation of laminar flow behavior.
Optimal Topology of Aircraft Rib and Spar Structures under Aeroelastic Loads
NASA Technical Reports Server (NTRS)
Stanford, Bret K.; Dunning, Peter D.
2014-01-01
Several topology optimization problems are conducted within the ribs and spars of a wing box. It is desired to locate the best position of lightening holes, truss/cross-bracing, etc. A variety of aeroelastic metrics are isolated for each of these problems: elastic wing compliance under trim loads and taxi loads, stress distribution, and crushing loads. Aileron effectiveness under a constant roll rate is considered, as are dynamic metrics: natural vibration frequency and flutter. This approach helps uncover the relationship between topology and aeroelasticity in subsonic transport wings, and can therefore aid in understanding the complex aircraft design process which must eventually consider all these metrics and load cases simultaneously.
Multidisciplinary Design Optimization on Conceptual Design of Aero-engine
NASA Astrophysics Data System (ADS)
Zhang, Xiao-bo; Wang, Zhan-xue; Zhou, Li; Liu, Zeng-wen
2016-06-01
In order to obtain better integrated performance of aero-engine during the conceptual design stage, multiple disciplines such as aerodynamics, structure, weight, and aircraft mission are required. Unfortunately, the couplings between these disciplines make it difficult to model or solve by conventional method. MDO (Multidisciplinary Design Optimization) methodology which can well deal with couplings of disciplines is considered to solve this coupled problem. Approximation method, optimization method, coordination method, and modeling method for MDO framework are deeply analyzed. For obtaining the more efficient MDO framework, an improved CSSO (Concurrent Subspace Optimization) strategy which is based on DOE (Design Of Experiment) and RSM (Response Surface Model) methods is proposed in this paper; and an improved DE (Differential Evolution) algorithm is recommended to solve the system-level and discipline-level optimization problems in MDO framework. The improved CSSO strategy and DE algorithm are evaluated by utilizing the numerical test problem. The result shows that the efficiency of improved methods proposed by this paper is significantly increased. The coupled problem of VCE (Variable Cycle Engine) conceptual design is solved by utilizing improved CSSO strategy, and the design parameter given by improved CSSO strategy is better than the original one. The integrated performance of VCE is significantly improved.
Design considerations for composite fuselage structure of commercial transport aircraft
NASA Technical Reports Server (NTRS)
Davis, G. W.; Sakata, I. F.
1981-01-01
The structural, manufacturing, and service and environmental considerations that could impact the design of composite fuselage structure for commercial transport aircraft application were explored. The severity of these considerations was assessed and the principal design drivers delineated. Technical issues and potential problem areas which must be resolved before sufficient confidence is established to commit to composite materials were defined. The key issues considered are: definition of composite fuselage design specifications, damage tolerance, and crashworthiness.
Cooling Panel Optimization for the Active Cooling System of a Hypersonic Aircraft
NASA Technical Reports Server (NTRS)
Youn, B.; Mills, A. F.
1995-01-01
Optimization of cooling panels for an active cooling system of a hypersonic aircraft is explored. The flow passages are of rectangular cross section with one wall heated. An analytical fin-type model for incompressible flow in smooth-wall rectangular ducts with coupled wall conduction is proposed. Based on this model, the a flow rate of coolant to each design minimum mass flow rate or coolant for a single cooling panel is obtained by satisfying hydrodynamic, thermal, and Mach number constraints. Also, the sensitivity of the optimal mass flow rate of coolant to each design variable is investigated. In addition, numerical solutions for constant property flow in rectangular ducts, with one side rib-roughened and coupled wall conduction, are obtained using a k-epsilon and wall function turbulence model, these results are compared with predictions of the analytical model.
Probabilistic Methods for Uncertainty Propagation Applied to Aircraft Design
NASA Technical Reports Server (NTRS)
Green, Lawrence L.; Lin, Hong-Zong; Khalessi, Mohammad R.
2002-01-01
Three methods of probabilistic uncertainty propagation and quantification (the method of moments, Monte Carlo simulation, and a nongradient simulation search method) are applied to an aircraft analysis and conceptual design program to demonstrate design under uncertainty. The chosen example problems appear to have discontinuous design spaces and thus these examples pose difficulties for many popular methods of uncertainty propagation and quantification. However, specific implementation features of the first and third methods chosen for use in this study enable successful propagation of small uncertainties through the program. Input uncertainties in two configuration design variables are considered. Uncertainties in aircraft weight are computed. The effects of specifying required levels of constraint satisfaction with specified levels of input uncertainty are also demonstrated. The results show, as expected, that the designs under uncertainty are typically heavier and more conservative than those in which no input uncertainties exist.
Design of a 4-seat, general aviation, electric aircraft
NASA Astrophysics Data System (ADS)
Rajagopalan, Arvindhakshan
Range and payload of current electric aircraft is limited primarily due to low energy density of batteries. However, recent advances in battery technology promise storage of more than 1 kWh of energy per kilogram of weight in the near future. This kind of energy storage makes possible the design of an electric aircraft comparable to, if not better than existing state-of-the art general aviation aircraft powered by internal combustion engines. This thesis explores through parametric studies the effect of lift-to-drag ratio, flight speed, and cruise altitude on required thrust power and battery energy and presents the conceptual and preliminary design of a four-seat, general aviation electric aircraft with a takeoff weight of 1750 kg, a range of 800 km, and a cruise speed of 200 km/h. An innovative configuration design will take full advantage of the electric propulsion system, while a Lithium-Polymer battery and a DC brush less motor will provide the power. Advanced aerodynamics will explore the greatest possible extend of laminar flow on the fuselage, the wing, and the empennage surfaces to minimize drag, while advanced composite structures will provide the greatest possible savings on empty weight. The proposed design is intended to be certifiable under current FAR 23 requirements.
Some design considerations for solar-powered aircraft
NASA Technical Reports Server (NTRS)
Phillips, W. H.
1980-01-01
Performance and operating characteristics are presented for a solar powered aircraft intended to remain aloft for long periods. The critical technologies which limit the performance are identified. By using the techniques presented, the effects of variation in the system parameters are studied. Practical design consideration are discussed.
Aircraft-crash-locating transmitter features design improvements
NASA Technical Reports Server (NTRS)
Manoli, R.; Ulrich, B. R.
1971-01-01
Crash locater is automatically ejected from aircraft at time of crash and begins transmitting at emergency radio frequencies monitored by all airports and airport control towers. Advantages are smaller size, simpler design and installation, extended transmitting range and life, greater deployment reliability and increased crash resistance.
Design Oriented Structural Modeling for Airplane Conceptual Design Optimization
NASA Technical Reports Server (NTRS)
Livne, Eli
1999-01-01
The main goal for research conducted with the support of this grant was to develop design oriented structural optimization methods for the conceptual design of airplanes. Traditionally in conceptual design airframe weight is estimated based on statistical equations developed over years of fitting airplane weight data in data bases of similar existing air- planes. Utilization of such regression equations for the design of new airplanes can be justified only if the new air-planes use structural technology similar to the technology on the airplanes in those weight data bases. If any new structural technology is to be pursued or any new unconventional configurations designed the statistical weight equations cannot be used. In such cases any structural weight estimation must be based on rigorous "physics based" structural analysis and optimization of the airframes under consideration. Work under this grant progressed to explore airframe design-oriented structural optimization techniques along two lines of research: methods based on "fast" design oriented finite element technology and methods based on equivalent plate / equivalent shell models of airframes, in which the vehicle is modelled as an assembly of plate and shell components, each simulating a lifting surface or nacelle / fuselage pieces. Since response to changes in geometry are essential in conceptual design of airplanes, as well as the capability to optimize the shape itself, research supported by this grant sought to develop efficient techniques for parametrization of airplane shape and sensitivity analysis with respect to shape design variables. Towards the end of the grant period a prototype automated structural analysis code designed to work with the NASA Aircraft Synthesis conceptual design code ACS= was delivered to NASA Ames.
NASA Technical Reports Server (NTRS)
2004-01-01
The grant closure report is organized in the following four chapters: Chapter describes the two research areas Design optimization and Solid mechanics. Ten journal publications are listed in the second chapter. Five highlights is the subject matter of chapter three. CHAPTER 1. The Design Optimization Test Bed CometBoards. CHAPTER 2. Solid Mechanics: Integrated Force Method of Analysis. CHAPTER 3. Five Highlights: Neural Network and Regression Methods Demonstrated in the Design Optimization of a Subsonic Aircraft. Neural Network and Regression Soft Model Extended for PX-300 Aircraft Engine. Engine with Regression and Neural Network Approximators Designed. Cascade Optimization Strategy with Neural network and Regression Approximations Demonstrated on a Preliminary Aircraft Engine Design. Neural Network and Regression Approximations Used in Aircraft Design.
NASA Technical Reports Server (NTRS)
Samareh, Jamshid A.; Sensmeier, mark D.; Stewart, Bret A.
2006-01-01
Algorithms for rapid generation of moderate-fidelity structural finite element models of air vehicle structures to allow more accurate weight estimation earlier in the vehicle design process have been developed. Application of these algorithms should help to rapidly assess many structural layouts before the start of the preliminary design phase and eliminate weight penalties imposed when actual structure weights exceed those estimated during conceptual design. By defining the structural topology in a fully parametric manner, the structure can be mapped to arbitrary vehicle configurations being considered during conceptual design optimization. Recent enhancements to this approach include the porting of the algorithms to a platform-independent software language Python, and modifications to specifically consider morphing aircraft-type configurations. Two sample cases which illustrate these recent developments are presented.
An aircraft model for the AIAA controls design challenge
NASA Technical Reports Server (NTRS)
Brumbaugh, Randal W.
1991-01-01
A generic, state-of-the-art, high-performance aircraft model, including detailed, full-envelope, nonlinear aerodynamics, and full-envelope thrust and first-order engine response data is described. While this model was primarily developed Controls Design Challenge, the availability of such a model provides a common focus for research in aeronautical control theory and methodology. An implementation of this model using the FORTRAN computer language, associated routines furnished with the aircraft model, and techniques for interfacing these routines to external procedures is also described. Figures showing vehicle geometry, surfaces, and sign conventions are included.
Visualization in the Design of Modern Aircraft Aerodynamics
NASA Technical Reports Server (NTRS)
Bryson, Steve; Kutler, Paul (Technical Monitor)
1997-01-01
Modem aircraft design involves study of airflow through both windtunnel testing and computer simulation. These computer simulations result in often very large and complex sets of numbers, which contain information critical to the aircrafts performance. This talk will describe how visualization is used to understand these simulations, using a variety of techniques including low-level analysis such as simulated particles, high-level feature detection, and virtual-reality-based techniques for exploration. We will focus on the challenges of extremely large data sets, interactive performance, and information extraction. The talk will close with a vision of the future including the integration of simulation and visualization.
Interactive Graphics Analysis for Aircraft Design
NASA Technical Reports Server (NTRS)
Townsend, J. C.
1983-01-01
Program uses higher-order far field drag minimization. Computer program WDES WDEM preliminary aerodynamic design tool for one or two interacting, subsonic lifting surfaces. Subcritical wing design code employs higher-order far-field drag minimization technique. Linearized aerodynamic theory used. Program written in FORTRAN IV.
Engine Conceptual Design Studies for a Hybrid Wing Body Aircraft
NASA Technical Reports Server (NTRS)
Tong, Michael T.; Jones, Scott M.; Haller, William J.; Handschuh, Robert F.
2009-01-01
Worldwide concerns of air quality and climate change have made environmental protection one of the most critical issues in aviation today. NASA's current Fundamental Aeronautics research program is directed at three generations of aircraft in the near, mid and far term, with initial operating capability around 2015, 2020, and 2030, respectively. Each generation has associated goals for fuel burn, NOx, noise, and field-length reductions relative to today's aircrafts. The research for the 2020 generation is directed at enabling a hybrid wing body (HWB) aircraft to meet NASA's aggressive technology goals. This paper presents the conceptual cycle and mechanical designs of the two engine concepts, podded and embedded systems, which were proposed for a HWB cargo freighter. They are expected to offer significant benefits in noise reductions without compromising the fuel burn.
Engine Conceptual Design Studies for a Hybrid Wing Body Aircraft
NASA Technical Reports Server (NTRS)
Tong, Michael T.; Jones, Scott M.; Haller, William J.; Handschuh, Robert F.
2009-01-01
Worldwide concerns of air quality and climate change have made environmental protection one of the most critical issues in aviation today. NASA s current Fundamental Aeronautics Research program is directed at three generations of aircraft in the near, mid and far term, with initial operating capability around 2015, 2020, and 2030, respectively. Each generation has associated goals for fuel burn, NOx, noise, and field-length reductions relative to today s aircrafts. The research for the 2020 generation is directed at enabling a hybrid wing body (HWB) aircraft to meet NASA s aggressive technology goals. This paper presents the conceptual cycle and mechanical designs of the two engine concepts, podded and embedded systems, which were proposed for a HWB cargo freighter. They are expected to offer significant benefits in noise reductions without compromising the fuel burn.
Neural Network and Regression Approximations Used in Aircraft Design
NASA Technical Reports Server (NTRS)
Patnaik, Surya N.; Hopkins, Dale A.; Lavelle, Thomas M.
1999-01-01
NASA Lewis Research Center's CometBoards Test Bed was used to create regression and neural network models for a High-Speed Civil Transport (HSCT) aircraft. Both approximation models that replaced the actual analysis tool predicted the aircraft response in a trivial computational effort. The models allow engineers to quickly study the effects of design variables on constraint and objective values for a given aircraft configuration. For example, an engineer can change the engine size by 1000 pounds of thrust and quickly see how this change affects all the output values without rerunning the entire simulation. In addition, an engineer can change a constraint and use the approximation models to quickly reoptimize the configuration. Generating the neural network and the regression models is a time-consuming process, but this exercise has to be carried out only once. Furthermore, an automated process can reduce calculations substantially.
NASA Technical Reports Server (NTRS)
Lax, F. M.
1975-01-01
A time-controlled navigation system applicable to the descent phase of flight for airline transport aircraft was developed and simulated. The design incorporates the linear discrete-time sampled-data version of the linearized continuous-time system describing the aircraft's aerodynamics. Using optimal linear quadratic control techniques, an optimal deterministic control regulator which is implementable on an airborne computer is designed. The navigation controller assists the pilot in complying with assigned times of arrival along a four-dimensional flight path in the presence of wind disturbances. The strategic air traffic control concept is also described, followed by the design of a strategic control descent path. A strategy for determining possible times of arrival at specified waypoints along the descent path and for generating the corresponding route-time profiles that are within the performance capabilities of the aircraft is presented. Using a mathematical model of the Boeing 707-320B aircraft along with a Boeing 707 cockpit simulator interfaced with an Adage AGT-30 digital computer, a real-time simulation of the complete aircraft aerodynamics was achieved. The strategic four-dimensional navigation controller for longitudinal dynamics was tested on the nonlinear aircraft model in the presence of 15, 30, and 45 knot head-winds. The results indicate that the controller preserved the desired accuracy and precision of a time-controlled aircraft navigation system.