Modified VISTA/F-16 piloted simulation study

A method of designing large envelope flight controllers for high performance aircraft is presented. The approach combines the simple two-degree of freedom architecture and recent advances in parameter-space control design techniques to form a new approach for large envelope flight control design. The procedure enables the designer to explicitly define the desired closed-loop dynamics and ensures that the closed-loop is stable and satisfies weighted frequency response magnitude (H/sub /spl infin//) specifications. The result is a straightforward procedure that enables the design of a robust flight controller that "forces" the closed-loop dynamics to behave like the specified 'desired dynamics' despite disturbances, modeling uncertainties, and variations in aircraft dynamics due to changing flight conditions. The procedure is presented by designing a pitch-rate controller for the F-16 Variable Stability In-Flight Simulator Test Aircraft (VISTA). Finally, the controller is compared to a 7/sup th/ order and a 10/sup th/ order linear parameter-varying controller and found to provide better performance and robustness.

This thesis outlines the effort in developing an integrated simulation framework for assessing the stability and control characteristics of an Unmanned Aerial Vehicle (UAV). The objectives are to understand flight instabilities experienced by the UAV and to propose solutions for the flight stability and control problems via an integrated modeling framework. The basic core of the framework involves the incorporation and integration of the UAV aerodynamic characteristics, flight stability derivatives and digital flight control system, into a six degree-of-freedom flight simulation model, the outputs of which are visualized graphically. To support the development of the simulation model, a comprehensive aerodynamic database for the UAV is required. Empirical aerodynamic estimation software such as USAF Digital Datcom and computational flow simulation software suites such as FLUENT and Solidworks Flo Works have been used to estimate the aerodynamic force and moments as well as stability and control derivatives of the UAV. By combining these estimates with wind tunnel data, a comprehensive aerodynamic characteristics model is constructed for augmenting the flight dynamics model of the UAV. Initially, a de-coupled longitudinal and lateral stability analysis based on the small disturbance theory has been developed to understand the stability modes of the UAV. A coupled longitudinal and lateral stability analysis was then developed to study the effects of the aerodynamic coupling caused by the single propulsion system to the stability and control characteristics. The model is created within the programmable software environments of d at lab^, ~imulink' and Aerospace s lock set@ which are used to construct the six-degree-of-freedom model for the dynamical motion of the UAV. The inputs to the model include the aerodynamic information based on the aerodynamic characteristics, stability derivatives, engine characteristics, flight variables and atmospheric disturbances. The outputs from this model are the responses of the flight variables depicting lateral, longitudinal and directional stability characteristics of the flight vehicle. Virtual Reality based systems such as the Flight Gear software has been used as a graphical interface for displaying the UAV trajectory, the orientation of the vehicle and the stability and control characteristics of the UAV based on the outputs from the model.

An exploratory investigation was conducted of the nonlinear aerodynamic and stability characteristics of a tailless generic fighter configuration featuring a chine-shaped forebody coupled to a slender cropped delta wing in the NASA Langley Research Center's 12-Foot Low-Speed Wind Tunnel. Forebody and wing vortex flow mechanisms were identified through off-body flow visualizations to explain the trends in the longitudinal and lateral-directional characteristics at extreme attitudes (angles of attack and sideslip). The interactions of the vortical motions with centerline and wing-mounted vertical tail surfaces were studied and the flow phenomena were correlated with the configuration forces and moments. Single degree of freedom, free-to-roll tests were used to study the wing rock susceptibility of the generic fighter model. Modifications to the nose region of the chine forebody were examined and fluid mechanisms were established to account for their ineffectiveness in modulating the highly interactive forebody and wing vortex systems.

This paper presents the survey of latest research on Model Reference Adaptive Control (MRAC). Effort is to present state of the art, find challenges, limitations and further scope of research on MRAC. Due to recent orientation towards soft computing research, publications with fuzzy logic and neural network have accelerated. Considering this point a section has been allocated to it. A section of variable structure MRAC has been included as variable structure control (VSC) is considered to be recent potential area. Many advanced techniques in MRAC have been surveyed and compiled in paper. Founded with focus on flight control and other aerospace applications, now MRAC has been designed and implemented for diverse application. Different applications of MRAC also have been reviewed here.

High-speed UAV flight control is challenged by unknown external aerodynamic disturbances and internal system variations due to complex aerodynamic configuration, propellant consumption, the center of gravity movement, and possible actuator failures. Redundant aerodynamic control effectors and engine inlet control mechanisms are exploited/blended to maintain adequate force/moment control to satisfy flight control reliability requirements. A control allocation (CA) mixing matrix is an essential element of the redundancy design in addressing fault tolerance capabilities. This paper employs an Intelligent Flight Controller (IFC) implemented in an adaptive control augmentation fashion to assist the UAV’s primary Nonlinear Dynamic Inversion (NDI) controller in restoring aircraft stability and command following objective when subjected to effector performance degradation, including failures. Through a blended design between optimal control modification (OCM) and derivative-free model reference adaptive control (DF-MRAC), the IFC offers performance consistency in comparison to traditional MRAC. The proposed IFC framework has demonstrated its effectiveness in assisting the baseline NDI flight control system to maintain its mission subjected to actuator’s failures (via an implicit automatic CA ‘re-distributing’ action). Furthermore, the IFC also works well with any existing onboard CA algorithm in dealing with effector failures without requiring restructuring of the CA blending matrix. It therefore deserves consideration for application to future high Mach UAVs.

Inspired by the first successful tests of a private manned spaceplane in 2004, the Student Aerospace Challenge was created in 2006 by the European Astronaut Club and its partners - Dassault Aviation, the European Space Agency, the International Astronautical Federation, Safran and Thales at the time - to allow European university students to explore some aspects of manned suborbital vehicles. Until 2020, the Challenge focused on a local reusable vehicle reaching Mach 3.5 and an altitude of 100 km. Since the 15th edition, to better respond to the evolution of the sector, a second vehicle is proposed: a hypersonic vehicle dedicated to point-to-point transportation taking, for example, less than two hours to travel from Barcelona to Tokyo. Each year, the Steering Committee defines several work packages corresponding to a large variety of study domains realistically related to this type of innovative vehicles like aerodynamic and flight control, structure, reusable propulsion, airworthiness, promotion, market analysis, legal frame & medicine. The introduction of a second vehicle having a quite different mission led the Committee to introduce dedicated topics. In addition, for the current edition, a new work package was proposed to cover potential applications of suborbital flights other than carrying passengers. In function of their background and interest, European University students have the opportunity to work, during several months, on a topic related to one of the work packages and to explore new solutions. Proposed projects should be technically realistic, economically viable and environmentally friendly. Reports and posters issued by student teams are evaluated by the Steering Committee some weeks before the “Suborbital Day”, a dedicated event organised like a mini-symposium, usually on-site where students present orally their projects and meet representatives of the different partners. The best-quoted projects are rewarded with prizes, among them, the ESA Grand Prize offering the winner team the unique opportunity to present their project in an appropriate European space-related event. To date, 216 teams and 998 University students coming from all over Europe already took part in the Student Aerospace Challenge, a motivating and ambitious multidisciplinary educational programme. Their participation allowed them to complement their knowledge, learn new skills and enlarge their network in the space sector

Thrust Vectoring Applied to Catastrophic Failure Prevention in Jet Transports

Wake vortex safety analysis and control for unmanned aircraft systems (UAS) operating alongside commercial aircraft in the terminal zones are critical aspects in the integration of UAS into the National Airspace System (NAS), particularly in terminal airspace. We address an approach that integrates reduced-order models of wake generation, evolution, interaction and its control incorporating both theoretical and high-fidelity predictions. Our previous study [1] estimated UAS responses in terms of the roll control ratios (RCR) developed by the follower aircraft depending on their position and orientation relative to the wake shed by the leader aircraft, as well as the duration of the wake-UAS interaction. The current paper complements and expands the previous study to describe a predictive, robust feedback-loop flight control model that is applicable to various classes of UAVs in unsteady flight-path scenarios including control of wake response. The control design presented in this paper extends previous research results by first demonstrating asymptotic (zero steady-state error) altitude regulation control in the presence of unmodeled vertical wind gust disturbances. To address the practical considerations involved in small UAV applications with limited computational resources, the proposed control method is designed with a computationally simplistic structure, without the requirement of complex calculations or function approximators in the control loop. Proof of the theoretical result is summarized, and detailed numerical simulation results are provided, which demonstrate the capability of the proposed nonlinear control method to asymptotically reject wind gust disturbances and parameter variations in the state space model. Simulation comparisons with a standard linear control method are provided for completeness.

Water tunnel facilities and flow visualization techniques have been developed at Northrop to provide high quality visualization of vortex interactions at high angles of attack. Results have provided considerable insight into highly complex three-dimensional flowfields generated by contemporary fighter aircraft. Studies have been made of the leading edge vortex systems generated by wing leading edge extensions (LEXs) typical of current subsonic-transonic fighter aircraft, high angle of attack aerodynamic asymmetries associated with the vortex system developed on long slender forebodies, and forebody and wing/LEX vortex interactions characteristic of highly maneuverable aircraft with hybrid wing planforms and slender forebodies. Qualitative results are in excellent agreement with existing subsonic wind tunnel data. The water-to-air analogy has been verified, that is, aircraft forebody and wing/LEX vortex systems, vortex system interactions, and the downstream influence on flow characteristics exhibited in air at high Reynolds numbers can be simulated at low Reynolds numbers in the water tunnel.

Improving the understanding of flight dynamics and control response characteristics of aircraft in upset conditions has been the focus of extensive research and development in recent years. Aircraft may experience upsets due to a number factors, and without suitable corrective actions, upsets may ultimately lead to loss of control in-flight (LOC-I) accidents. To provide models for engineering analysis and pilot training, conventional simulation approaches have historically used multi-variable aerodynamic databases (look-up tables) derived from wind tunnel and/or flight test data, which are difficult to acquire and do not always encompass the desired flight envelope. An alternative approach for modeling aerodynamic and flight dynamic characteristics of aircraft in upset conditions has been investigated based on methods that are not data-driven but that are derived from physicallybased models. This paper describes a modeling approach based on nonlinear lifting line/strip theory analysis of the aerodynamic loads combined with an unsteady, freelyevolving vortex wake model for the induced velocity due to the lifting surfaces. The paper describes background work and recent developments toward a generic, physically-based model for aircraft upsets. Application of the modeling approach to unmanned aircraft is presented.

Aircraft with a twin vertical tail and leading-edge extension configuration may experience vertical tail buffeting during high-angle-of-attack maneuvering flight. This issue can lead to structural fatigue damage in the vertical tail, shortening its service life and increasing maintenance costs, ultimately compromising flight safety. Therefore, accurate prediction of buffeting loads and responses is essential during design. In the preliminary stage, wind tunnel testing is the primary means to obtain dynamic data such as fluctuating pressure and acceleration response, which can be transformed to full-scale conditions through similitude principles. However, the elastic scaling model used in buffeting tests is usually established for a specific flight condition. When the flow velocity or objective flight condition changes, the similitude relationship becomes invalid, limiting the applicability of test results and preventing full-envelope strength verification. To overcome this limitation, this study proposes a multi-objective point response prediction method for vertical tail buffeting. The method enables the prediction of full-scale responses at multiple objective flight conditions using wind tunnel data that do not strictly satisfy similitude criteria. A complete aircraft vertical tail buffet (rigid/elastic) hybrid model was developed for testing, and an Adjusted Model incorporating elastic scaling transformation was established. The proposed method was validated through experiments, demonstrating improved test data utilization and prediction accuracy across multiple-objective flight conditions.

The NASA Flying Controls Testbed (FLiC) is a relatively small and inexpensive unmanned aerial vehicle developed specifically to test highly experimental flight control approaches. The most recent version of the FLiC is configured with 16 independent aileron segments, supports the implementation of C-coded experimental controllers, and is capable of fully autonomous flight from takeoff roll to landing, including flight test maneuvers. The test vehicle is basically a modified Army target drone, AN/FQM-117B, developed as part of a collaboration between the Aviation Applied Technology Directorate (AATD) at Fort Eustis, Virginia and NASA Langley Research Center. Several vehicles have been constructed and collectively have flown over 600 successful test flights, including a fully autonomous demonstration at the Association of Unmanned Vehicle Systems International (AUVSI) UAV Demo 2005. Simulations based on wind tunnel data are being used to further develop advanced controllers for implementation and flight test.

This paper addresses flight control system synthesis and the accommodation of controlled plant variation in an aircraft design envelope by using the frequency domain based quantitative feedback theory robust control system design method. The plant variations considered include varying flight conditions in the flight envelope and damage to aerodynamic control surfaces. A digital flight controller for the aircraft's longitudinal channel is designed and validated with simulations containing nonlinear saturation elements.

Airbus and Boeing are cooperatively presenting this topic dealing with transport aircraft stalls. The paper will begin by defining a stall, followed by a review of requirements, predictive validation and flight testing. There are various ways of designing modern jet transports for the stall regime such as aerodynamic approaches, flight deck indications, and augmentation control laws to deal with the high angle-of-attack (α) arena. The goal of augmented control laws for high α is common – no full aerodynamic stall or loss of climb performance should occur in the operational flight envelope, in Normal flight control modes. The validation techniques employed in preparation for a flight test campaign will follow. These include flight characteristic predictions based on wind-tunnel data as well as pilot-in-the-loop simulation rehearsals. The preparation for flight testing will be reviewed from both the engineer and pilot viewpoints. This will be followed by a review of various flight testing that has been conducted. The paper will close with a brief foray into what the future of transport stalls could be – perhaps protection features in degraded flight control modes? What are the benefits as well as drawbacks to increased augmentation for high α?

A key consideration in the development of flight control systems early in the design stage is the availability of aerodynamic information for different candidate configurations. Aerodynamic estimation methods must be available to provide the connection between the configuration geometry and its stability and control characteristics. The accuracy of current estimation methods is studied in this paper through application of DATCOM and APAS to available XB-70 wind tunnel and flight test data. The study was carried out for the subsonic approach condition and three supersonic conditions. Tables and charts are presented to provide a quantitative assessment of the accuracy of the predictions, and areas requiring improvement are identified. Results show that APAS and DATCOM predictions are good for most lateral/directional stability and control derivatives. Estimations for the pitching moment slope, yaw damping and yawing moment due to flap deflection derivatives are only fair. The most difficult derivative to predict is the rolling moment due to sideslip.