NASA STAR REPORTS_ 09_09_05 - Aircraft Stability and Control

Clarke, Robert; Allen, Michael J.; Dibley, Ryan P.; Gera, Joseph; Hodgkinson, John; [2005]; 31 pp.; In English; AIAA Atmospheric Flight Mechanics Conference, 15-18 Aug.2005, San Francisco, CA, USA; No Copyright; Avail: CASI; A03, Hardcopy

Successful flight-testing of the Active Aeroelastic Wing airplane was completed in March 2005. This program, which started in 1996, was a joint activity sponsored by NASA, Air Force Research Laboratory, and industry contractors. The test program contained two flight test phases conducted in early 2003 and early 2005. During the first phase of flight test, aerodynamic models and load models of the wing control surfaces and wing structure were developed. Design teams built new research control laws for the Active Aeroelastic Wing airplane using these flight-validated models; and throughout the final phase of flight test, these new control laws were demonstrated. The control laws were designed to optimize strategies for moving the wing control surfaces to maximize roll rates in the transonic and supersonic flight regimes. Control surface hinge moments and wing loads were constrained to remain within hydraulic and load limits. This paper describes briefly the flight control system architecture as well as the design approach used by Active AeroelasticWing project engineers to develop flight control system gains. Additionally, this paper presents flight test techniques and comparison between flight test results and predictions. Author

Flight Tests; Aeroelastic Research Wings; Control Systems Design; Flight Control

20050206089 Sytronics, Inc., Dayton, OH USA

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Exploring Automation Issues in Supervisory Control of Multiple UAVs

Ruff, Heath A.; Calhoun, Gloria L.; Draper, Mark H.; Fontejon, John V.; Guilfoos, Brian J.; Mar. 2004; 7 pp.; In English; Original contains color illustrations Report No.(s): AD-A436181; No Copyright; Avail: CASI; A02, Hardcopy

An evaluation was conducted on a generic UAV operator interface simulation testbed to explore the effects of levels-of-automation (LOAs) and automation reliability on the number of simulated UAVs that could be supervised by a single operator. LOAs included Management-by-Consent (operator consent required) and Management-by-Exception (action automatic unless operator declines). Results indicated that the tasks were manageable, but performance decreased with increased number of UAVs supervised and reduced automation reliability. Performance with the two LOAs varied little and did not show a consistent trend across measures. Analyses indicated that participants typically did not utilize the automation. A follow-on study was conducted that employed shorter LOA time limits. Results showed participants’ workload and confidence ratings were less favorable for the shorter limits and they still exercised the automation rarely, although more frequently. Further research is needed to explore the complex relationship between LOAs, time limits, perception of workload, vigilance effects, and confidence. DTIC

Flight Control; Human-Computer Interface; Pilotless Aircraft; Remotely Piloted Vehicles

20050206090 Barron Associates, Inc., Charlottesville, VA USA

Real-Time Parameter Identification for Self-Designing Flight Control

Ward, D. G.; Barron, R. L.; Carley, M. P.; Curtis, T. J.; Jan. 2005; 7 pp.; In English Contract(s)/Grant(s): F49620-93-C-0014 Report No.(s): AD-A436182; No Copyright; Avail: Defense Technical Information Center (DTIC)

A self-designing flight control system (SDFCS) could provide a cost-effective means for developing controllers for new aircraft by eliminating analyst-intensive design of numerous individual controllers, each optimized for a single flight condition. Additionally, the SDFCS could improve the capabilities of existing aircraft by enhancing control performance in new flight regimes such as high angle-of-attack or post-stall maneuvers. Finally, the SDFCS could automatically reconfigure the control system to account for sudden changes such as may result from airframe and/or effector impairment(s). Rapid identification of time-varying, nonlinear plants is an important enabling technology for most SDFCS concepts. In this paper, the authors present a modified sequential least squares (MSLS) parameter identification method and compare its performance to that of standard RLS techniques using a simulated nonlinear F-16 with multiaxes thrust-vectoring (MATV) aircraft. It is shown that MSLS offers significant improvement in performance over conventional RLS parameter identification by providing: (1) a recursive estimation algorithm that penalizes noisy estimates and is less subject to ill-conditioning as its forgetting factor is reduced; (2) detection of airframe and effector impairments and corresponding adjustments of the algorithm settings; and (3) an intelligent supervisor that injects a minimum level of effector random activity to ensure identifiability. DTIC

Flight Control; Independent Variables; Parameter Identification; Real Time Operation

20050206150 Optimal Synthesis, Inc., Palo Alto, CA USA

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Computer-Aided Synthesis of Nonlinear Autopilots for Missiles

Menon, P. K.; Ohlmeyer, E. J.; Jan. 2004; 27 pp.; In English Contract(s)/Grant(s): N00024-97-C-4178 Report No.(s): AD-A436320; No Copyright; Avail: Defense Technical Information Center (DTIC)

Powerful nonlinear approaches for missile autopilot design have recently emerged in the literature, which have the potential to deliver improved missile performance. However, the lack of computational methods has made it difficult for the practicing engineers to exploit these techniques in routine applications. Another factor that has slowed their application is that the missile models are generally available in the form of simulations, rather than as compact set of differential-algebraic equations. This paper discusses five different approaches for computer-aided nonlinear control system design that ameliorate these difficulties. Since these design techniques are based on simulation models, they enable direct synthesis of nonlinear autopilots using missile models of arbitrary complexity. Airframe stabilization of a nonlinear, longitudinal missile model is used to illustrate the design techniques. DTIC

Automatic Pilots; Computer Aided Design; Computer Techniques; Missiles; Nonlinearity

20050206152 Optimal Synthesis, Inc., Palo Alto, CA USA

Optimal Fixed-Interval Integrated Guidance-Control Laws for Hit-to-Kill Missiles

Menon, P. K.; Sweriduk, G. D.; Ohlmeyer, E. J.; Aug. 2003; 10 pp.; In English; Original contains color illustrations Contract(s)/Grant(s): N00178-01-C-1020 Report No.(s): AD-A436322; AIAA-2003-5579; No Copyright; Avail: Defense Technical Information Center (DTIC)

Due to their potential for reducing the weapon size and efficiency, design methods for realizing hit-to- kill capabilities in missile systems are of significant research interest in the missile flight control community. As defined in this paper, hit-to-kill capability requires the missile to consistently achieve point-mass miss distances less than half the minimum dimension of the target. It has been noted in the literature that the chief contributors to the miss distance in homing missiles are the seeker errors, autopilot lag, target maneuvers, and target state estimation lag. Guidance laws for ameliorating the effects of each of these miss distance components have been discussed in several recent publications. The present research addresses the hit-to-kill missile flight control problem by casting it as an integrated guidance-control problem. By including the complete dynamics of the missile, the integrated guidance-control formulation automatically compensates for the impact of the autopilot lag on the miss distance. The resulting finite-interval control problem is then solved using a transformation approach. Interception by a kinetic warhead is used as an example to illustrate the performance of the integrated guidance-control law. DTIC

Control Theory; Flight Control; Missiles

20050207434 NASA Langley Research Center, Hampton, VA, USA

Computational Methods for Stability and Control (COMSAC): The Time Has Come

Hall, Robert M.; Biedron, Robert T.; Ball, Douglas N.; Bogue, David R.; Chung, James; Green, Bradford E.; Grismer, Matthew J.; Brooks, Gregory P.; Chambers, Joseph R.; [2005]; 40 pp.; In English; AIAA Atmospheric Flight Mechanics Conference and Exhibit, 15-18 Aug. 2005, San Francisco, CA, USA Contract(s)/Grant(s): 23-065-30-11 Report No.(s): AIAA Paper 2005-6121; Copyright; Avail: CASI; A03, Hardcopy

Powerful computational fluid dynamics (CFD) tools have emerged that appear to offer significant benefits as an adjunct to the experimental methods used by the stability and control community to predict aerodynamic parameters. The decreasing costs for and increasing availability of computing hours are making these applications increasingly viable as time goes on and the cost of computing continues to drop. This paper summarizes the efforts of four organizations to utilize high-end computational fluid dynamics (CFD) tools to address the challenges of the stability and control arena. General motivation and the backdrop for these efforts will be summarized as well as examples of current applications. Author

Computational Fluid Dynamics; Aerodynamics; Aircraft Stability; Aircraft Control

20050207439 NASA Langley Research Center, Hampton, VA, USA

Post-Stall Aerodynamic Modeling and Gain-Scheduled Control Design

Wu, Fen; Gopalarathnam, Ashok; Kim, Sungwan; [2005]; 15 pp.; In English; AIAA Guidance, Navigation, and Control Conference and Exhibit, 15-18 Aug. 2005, San Francisco, CA, USA Contract(s)/Grant(s): 23-760-40-01 Report No.(s): AIAA Paper 2005-6169; Copyright; Avail: CASI; A03, Hardcopy

A multidisciplinary research e.ort that combines aerodynamic modeling and gain-scheduled control design for aircraft flight at post-stall conditions is described. The aerodynamic modeling uses a decambering approach for rapid prediction of post-stall aerodynamic characteristics of multiple-wing con.gurations using known section data. The approach is successful in bringing to light multiple solutions at post-stall angles of attack right during the iteration process. The predictions agree fairly well with experimental results from wind tunnel tests. The control research was focused on actuator saturation and .ight transition between low and high angles of attack regions for near- and post-stall aircraft using advanced LPV control techniques. The new control approaches maintain adequate control capability to handle high angle of attack aircraft control with stability and performance guarantee. Author

Multidisciplinary Research; Aircraft Control; Angle of Attack; Aerodynamic Characteristics; Control Systems Design; Stability

20050209962 NASA Langley Research Center, Hampton, VA, USA

Wake Vortex Tracking Using a 35 GHz Pulsed Doppler Radar

Neece, Robert T.; Britt, Charles L.; White, Joseph H.; Mudukutore, Ashok; Nguyen, Chi; Hooper, Bill; January 2005; 11 pp.; In English; 5th NASA Integrated Communications, Navigation, and Surveillance (ICNS) Conference andWorkshop, 2-5 May 2005, Fairfax, VA, USA; Copyright; Avail: CASI; A03, Hardcopy

A 35 GHz, pulsed-Doppler radar system has been designed and assembled for wake vortex detection and tracking in low visibility conditions.

Aircraft wake vortices continue to be an important factor in determining safe following distances or spacings for aircraft in the terminal area. Currently, under instrument meteorological conditions (IMC), aircraft adhere to conservative, fixed following-distance guidelines based primarily on aircraft weight classifications. When ambient conditions are such that vortices will either drift or dissipate, leaving the flight corridor clear, the prescribed spacings are unnecessarily long and result in decreased airport throughput. There is a potential for significant airport efficiency improvement, if a system can be employed to aid regulators and pilots in setting safe and efficient following distances based on airport conditions. The National Aeronautics and Space Administration (NASA), the Federal Aviation Agency, and Volpe National Transportation Systems Center have promoted and worked to develop systems that would increase airport capacity and provide for safe reductions in aircraft separation.

The NASA Aircraft Vortex Spacing System (AVOSS), a wake vortex spacing system that can provide dynamic adjustment of spacings based on real-time airport weather conditions, has demonstrated that Lidar systems can be successfully used to detect and track vortices in clear air conditions. To fill the need for detection capability in low-visibility conditions, a 35 GHz, pulsed-Doppler radar system is being investigated for use as a complimentary, low-visibility sensor for wake vortices. The radar sensor provides spatial and temporal information similar to that provided by Lidar, but under weather conditions that a Lidar cannot penetrate. Currently, we are analyzing the radar design based upon the data and experience gained during the wake vortex Lidar deployment with AVOSS at Dallas/Fort Worth International Airport. As part of this study, two numerical models were utilized in system simulations.

The results of this study improve our understanding of the method of detection, resolution requirements for range and azimuth, pulse compression, and performance prediction. Simulations applying pulse compression techniques show that detection is good in heavy fog to greater than 2000 m. Both compressed and uncompressed short pulses show the vortex structure. To explore operational challenges, siting and scanning strategies were also analyzed. Simulation results indicate that excellent wake vortex detection, tracking and classification is possible in drizzle (+15 dBZ) and heavy fog (- 13 dBZ) using short pulse techniques (\h99ns) at ranges on the order of 900 m, with a modest power of 500 W output. At 1600 m, detection can be expected at reflectivities as low as -13 dBZ (heavy fog).

The radar system, as designed and built, has the potential to support field studies of a wake vortex spacing system in low-visibility conditions ranging from heavy fog to rain, when sited within 2000m of the flight path. Author

Vortices; Wakes; Tracking (Position); Doppler Radar; Low Visibility

20050209974 National Aerospace Lab., Amsterdam, Netherlands

Aeroelasticity, Chapter 14

Meijer, J. J.; Introduction to Flight Test Engineering, Volume 14; July 2005, pp. 14-1 - 14-15; In English; See also 20050209967; Copyright; Avail: CASI; A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document

Flutter is an aeroelastic phenomenon in which aerodynamic, elastic and inertia forces couple unfavorably at a sufficiently high speed to produce an unstable oscillation which may grow without limit and so result in a structural failure. Flutter is, unfortunately, not a problem that will just ‘go away’: modern aircraft, in particular, are progressively more flexible, fly faster, and are highly control configured, more maneuverable and more system dependent. The development of active control systems for flutter suppression will actually complicate the problem by introducing a fourth ‘servo’ dimension into the aeroelastic scene and will at best only defer flutter to higher speeds. Derived from text

Aeroelasticity; Flutter; Vibration Damping; Control Systems Design; Aerodynamic Forces

20050209998 Fokker B.V., Schipol-Oost, Netherlands

Flight Envelope, Chapter 12

Walgemoed, H.; Introduction to Flight Test Engineering, Volume 14; July 2005, pp. 12-1 - 12-13; In English; See also 20050209967; Copyright; Avail: CASI; A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document

The term ‘flight envelope’ is used to refer to the boundaries of aircraft loading and flight conditions within which operation of the aircraft is satisfactory, and beyond which some aspect becomes unacceptable. This flight envelope represents, in fact, the limiting conditions arising from a matrix of inter-related flight envelopes covering the appropriate variables. Thus, for each loading (i.e., external stores configuration and its associated range of weight and center of gravity (c.g.) position) and aircraft configuration (i.e., position of undercarriage (u/c), flaps, slats, etc.), the envelopes of airspeed versus altitude, airspeed versus load factor, angle of attack versus angle of sideslip, etc., must be investigated to establish the limits within which all aspects such as handling qualities, engine behavior, structural loads, etc., remain acceptable.

Flight testing of new or derivative aircraft models is carried out with the initial purpose of defining a flight envelope which is, first and foremost, safe and secondarily, which enables the effective use of the vehicle for its intended purpose. Flight testing occurs only after numerous reviews of the design and review of results from ground tests and predictions of flight characteristics in such areas as structures, aerodynamics, stability and control, flight controls (particularly fly-by-wire control systems, propulsion, etc.). Accordingly, opening and expanding the envelope is a task that must be approached cautiously, systematically, and with coordination and cooperation of the many disciplines involved in the design and test of an airplane. (Sections 8 and 10 cover test planning and safety of flight considerations, respectively).

The fundamental tenet in establishing a flight envelope via flight test is risk reduction. This is reflected in the typical sequence of events leading to initial flight test - design reviews (both hardware and software), then ground test involving singular disciplines (windtunnel tests for aerodynamics, structurally loading the wing/fuselage/nacelle on a ground test article with loads anticipated to occur in flight, flight control system control law checkout, propulsion test cell runs and/or flying test bed tests, etc.), and then ground tests involving multi-disciplines (See Section 9). Only after these have been accomplished will an initial, limited, low-risk, flight envelope be established.

The limited envelope will typically be in the middle of the projected final flight envelope. Subsequent flight tests will then be devoted to expanding the initial envelope by operating the airplane at increasing ranges - representing increasing risk - of engine operation, airspeeds both fast and slow, altitude, load factor both above and below 1g, centers of gravity (fore and aft), and with system/subsystem failures. Whether flight tests are to define a flight envelope on a new model airplane with the attendant new airframe, new engine(s), and new subsystems (hydraulics, pressurization, etc.), or on an airplane involving only a few of these areas such as new engines in an old airframe, the fundamental approach to establishing an envelope is the same. Derived from text

Flight Envelopes; Flight Control; Flight Characteristics; Aerodynamic Stability; Aerodynamic Loads; Aircraft Models; Control Systems Design; Control Theory; Flight Tests

20050210000 Air Force Flight Test Center, Edwards AFB, CA, USA

Handling Qualities, Chapter 15

Lee, Robert E., Jr.; Introduction to Flight Test Engineering, Volume 14; July 2005, pp. 15-1 - 15-16; In English; See also 20050209967; Copyright; Avail: CASI; A03, Hardcopy; Available from CASI on CD-ROM only as part of the entire parent document

This section will set forth the types of tests, procedures, instrumentation requirements, data analysis and presentation, and the purposes for conducting the various tests required to define an aircraft’s handling qualities. In general, handling qualities tests are conducted to define the basic stability and control characteristics of the aircraft, define how the aircraft responds to pilot or other inputs, and to define techniques to maintain safe flight in the operational envelope. The terms handling qualities and flying qualities are often used synonymously; however, a distinction is sometimes made between the two notions. Flying qualities generally refers to a set of attributes determined by an aircraft’s stability and control characteristics. Handling qualities reflect the ease with which a pilot can carry out a mission task using an aircraft that possesses a particular set of flying qualities. There are a number of sources of detailed information regarding test conduct, analysis, presentation, and evaluation of handling qualities. A few of these sources are listed as references and several more are shown as bibliographic entries. Derived from text

Aircraft Control; Controllability; Flight Characteristics; Flight Envelopes; Stability

20050210003 NASA Langley Research Center, Hampton, VA, USA

An Investigation of the Impact of Aerodynamic Model Fidelity on Close-In Combat Effectiveness Prediction in Piloted Simulation

Persing, T. Ray; Bellish, Christine A.; Brandon, Jay; Kenney, P. Sean; Carzoo, Susan; Buttrill, Catherine; Guenther, Arlene; [2005]; 13 pp.; In English; AIAA Modeling and Simulation Technologies Conference and Exhibit, 15-18 Aug. 2005, San Francisco, CA, USA Contract(s)/Grant(s): 23R-066-50 Report No.(s): AIAA Paper 2005-5885; Copyright; Avail: CASI; A03, Hardcopy

Several aircraft airframe modeling approaches are currently being used in the DoD community for acquisition, threat evaluation, training, and other purposes. To date there has been no clear empirical study of the impact of airframe simulation fidelity on piloted real-time aircraft simulation study results, or when use of a particular level of fidelity is indicated. This paper documents a series of piloted simulation studies using three different levels of airframe model fidelity. This study was conducted using the NASA Langley Differential Maneuvering Simulator. Evaluations were conducted with three pilots for scenarios requiring extensive maneuvering of the airplanes during air combat. In many cases, a low-fidelity modified point-mass model may be sufficient to evaluate the combat effectiveness of the aircraft. However, in cases where high angle-of-attack flying qualities and aerodynamic performance are a factor or when precision tracking ability of the aircraft must be represented, use of high-fidelity models is indicated. Author

Aerodynamic Characteristics; Computerized Simulation; Models; Fighter Aircraft

Source: NASA.

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