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Robust dynamic inversion controller design and analysis for the X-38
Abstract
A new way to approach robust Dynamic Inversion controller synthesis is addressed in this paper. A Linear Quadratic Gaussian outer-loop controller improves the robustness of a Dynamic Inversion inner-loop controller in the presence of uncertainties. Desired dynamics are given by the dynamic compensator, which shapes the loop. The selected dynamics are based on both performance and stability robustness requirements. These requirements are straightforwardly formulated as frequency-dependent singular value bounds during synthesis of the controller. Performance and robustness of the designed controller is tested using a worst case time domain quadratic index, which is a simple but effective way to measure robustness due to parameter variation. Using this approach, a lateral-directional controller for the X-38 vehicle is designed and its robustness to parameter variations and disturbances is analyzed. The controller analysis is extended to the nonlinear system where both control input displacements and rates are bounded. Overall, this combination of controller synthesis and robustness criteria compares well with the usynthesis technique. It also is readily accessible to the practicing engineer, in terms of understanding and use.
... As is well known, another main weakness of NDI is the requirement of the plant's input distribution matrix to be invertible. However, it is often the case that the input distribution matrix becomes "almost" singular for some states and hence may lead to excessively large control activities [33]. As deflection range and rate limitations are inevitably present in physical actuated surfaces, aggressively piloted maneuvers of aircraft may yield actuator saturation and associated degradation of closed-loop performance or even loss of stability [34]. ...
... Note that the anti-windup modified disturbance observer is driven by the actual surface deflections rather than the surface commands. Combining the plant dynamics of Equation (42), the inner-loop NDI controller of Equation (33) and the anti-windup modified disturbance observer of Equation (48), it is readily reached that the closed-loop tracking error takes the same form as Equation (35). Hence, the effect of actuator saturation on the closed-loop tracking performance is also removed. ...
... Such an outcome was in agreement with the prospective effect of the disturbance observer. On one hand, the NDI-DO and the NDI-AW controllers estimated the lumped disturbances using Equations (32) and (48), respectively, which were then compensated using the inner-loop dynamic inversion controller of Equation (33). As a result, the influence of the lumped disturbances to the closed-loop system was mostly rejected by the disturbance observer augmented controllers. ...
This paper focuses on the design of a disturbance rejection controller for a tailless aircraft based on the technique of nonlinear dynamic inversion (NDI). The tailless aircraft model mounted on a three degree-of-freedom (3-DOF) dynamic rig in the wind tunnel is modeled as a nonlinear affine system subject to mismatched disturbances. First of all, a baseline NDI attitude controller is designed for sufficient stability and good reference tracking performance of the nominal system. Then, a nonlinear disturbance observer (NDO) is supplemented to the baseline NDI controller to estimate the lumped disturbances for compensation, including unmodeled dynamics, parameter uncertainties, and external disturbances. Mathematical analysis demonstrates the convergence of the employed NDO and the resulting closed-loop system. Furthermore, an anti-windup modification is applied to the NDO for control performance preserving in the presence of actuator saturation. Subsequently, the designed control schemes are preliminarily validated and compared via simulations. The baseline NDI controller demonstrates satisfactory attitude tracking performance in the case of nominal simulation; the NDO augmented NDI controller presents significantly improved ability of disturbance rejection when compared with the baseline NDI controller in the case of robust simulation; the anti-windup modified scheme, rather than the baseline NDI controller nor the NDO augmented NDI controller, can preserve the closed-loop performance in the case of actuator saturation. Finally, the baseline NDI scheme and the NDO augmented NDI scheme are implemented and further validated in the wind tunnel flight tests, which demonstrate that the experimental results are in good agreement with that of the simulations.
... In recent years, wind power generation technology has received extensive attention and attention from scholars in various countries. The wind farm is the control center for wind power generation [25,26] The wind power control method is the key issue of the current wind power generation technology [27], the more commonly used wind power control method has a strong robust control method [28], neural network fuzzy control [29], the best search for superior control method and so on. A fuzzy control method [27,30] based on the safety margin of backpressure protection for direct air-cooled units is presented. ...
... The wind farm is the control center for wind power generation [25,26] The wind power control method is the key issue of the current wind power generation technology [27], the more commonly used wind power control method has a strong robust control method [28], neural network fuzzy control [29], the best search for superior control method and so on. A fuzzy control method [27,30] based on the safety margin of backpressure protection for direct air-cooled units is presented. A new type of fuzzy controller is designed, and self-adjusting factor is introduced to adjust the fuzzy controller to suit the complex and changeable control system. ...
... The fan unit is mainly composed of wind turbine and asynchronous wind turbine. According to the mechanical characteristics of wind turbine, the output power of wind turbine is related to the wind speed [27], namely ...
... Then, the emphasis is to design an appropriate attitude control system to track the guidance command, which is the focus of this paper. Several approaches have been proposed in the past, such as gain scheduling (GS) [2,3], dynamic inversion (DI) technique [4][5][6], trajectory linearization control (TLC) [7][8][9], sliding mode control (SMC) [10,11], and state-dependent Riccati equation (SDRE) control [12,13]. They provided good control performance, but input constraint was not taken into account in the above research. ...
... Following the backstepping design procedure, the COM (5)-(6) is decomposed into attitude angle subsystem and attitude angular rate subsystem. The control scheme design begins from attitude angle subsystem (5) where the virtual control input is developed and used as the reference command for the angular rate subsystem (6). Then, the actual control input is designed. ...
... The bound of control authority is assumed to be max = 6 × 10 5 . Therefore, the control limits used for exploring the capability of the designed control strategy in adhering to the input constraint are u max = (6 × 10 5 6 × 10 5 uncertainty is assumed to be ΔI = 0.2I. The initial condition of reentry flight is given in Table 1. ...
This paper presents the finite time attitude tracking control problem of reusable launch vehicle (RLV) in reentry phase under input constraint, model uncertainty, and external disturbance. A control-oriented model of rotational dynamics is developed and used for controller design for the complex coupling of the translational and rotational dynamics. Firstly, fast terminal sliding mode control is incorporated into backstepping control to design controller considering model uncertainty and external disturbance. The “explosion of terms” problem inherent in backstepping control is eliminated by a robust second order filter. Secondly, the control problem in the presence of input constraint is further considered, and a constrained adaptive backstepping fast terminal sliding mode control scheme is developed. At the control design level, adaptive law is employed to estimate the unknown norm bound of lumped uncertainty with the reduction of computational burden. The Lyapunov-based stability analysis of the closed-loop system is carried out. The control performance is presented via the simulation of six-degree-of-freedom (6-DOF) model of RLV.
... However, they still require the priori model of the UAV and are sensitive to model uncertainties. Robust µ synthesis or linear quadratic Gaussian controllers [17], as well as model reference adaptive control (MRAC), are applied to reduce the model dependence [18]. However, the performance of the robust controller is conservative and it is hard to tune the high-order controller when applied on the real tilt-rotor UAV. ...
... According to the daisy-chain algorithm, a u and vr u are calculated sequentially in the two allocation steps. Firstly, the deflections of the aerodynamic effectors are derived: (17) where T c is the rotor's thrust command, ...
This paper proposes a unified attitude controller based on the modified linear active disturbance rejection control (LADRC) for a dual-tiltrotor unmanned aerial vehicle (UAV) with cyclic pitch to achieve accurate attitude control despite its nonlinear and time-varying characteristics during flight mode transitions. The proposed control algorithm has higher robustness against model mismatch compared with the model-based control algorithms. The modified LADRC utilizes the state feedbacks from the onboard sensors like IMU and Pitot tube instead of the mathematical model of the plane. It has less dependency on the accurate dynamics model of the dual-tiltrotor UAV, which can hardly be built. In contrast to the original LADRC, an actuator model is integrated into the modified LADRC to compensate for the non-negligible slow rotor flapping dynamics and servo dynamics. This modification eliminates the oscillation of the original LADRC when applied on the plant with slow-response actuators, such as propeller and rotors of the helicopter. In this way, the stability and performance of the controller are improved. The controller replaces the gain-scheduling or the control logic switching by a unified controller structure, which simplifies the design approach of the controller for different flight modes. The effectiveness of the modified LADRC and the performance of the unified attitude controller are demonstrated in both simulation and flight tests using a dual-tiltrotor UAV. The attitude control error is less than ±4° during the conversion flight. The control rising time in different flight modes is all about 0.5 s, despite the variations in the airspeed and tilt angle. The flight results show that the controller guarantees high control accuracy and uniform control quality in different flight modes.
... Consequently, several nonlinear attitude control methods have been proposed with the objective of achieving enhanced performance. A robust dynamic inversion controller design is discussed in [39] which has a linear quadratic Guassian (LQG)-based outer loop controller along with dynamic inversion-based inner loop controller. However, exact knowledge of the system dynamics is essential for the assured performance of a dynamic inversion controller. ...
... The numerical simulations on the 3-DoF dynamic model of the re-entry RLV are carried out with respect to the re-entry aerodynamic data in [39]. ...
In this work, a novel strategy for the inverse optimal control of a class of affine nonlinear systems is proposed. The proposed strategy involves translation of the infinite horizon nonlinear optimal control problem into a Diagonal stability problem, followed by the formulation of a set of criteria for the synthesis of a stabilizing feedback control law. Hence, the proposed controller design methodology is refered to as inverse optimal control via diagonal stabilization. Besides providing a closed-form solution, the methodology also possesses the added advantage of inherent robustness, on account of adequate stability margins. Most importantly, the methodology ensures an estimate of the associated domain of attraction, which is a highly desirable feature especially in the case of crucial and stringent aerospace applications. The proposed methodology is applied for the re-entry control of a reusable launch vehicle which provides a full envelope optimality-based design philosophy. For this, the control-oriented attitude model based on the three-degree-of-freedom dynamic model is utilized. The resulting control law possess desirable features of guaranteed estimate of stability domain, assured design flexibility, and inherent robustness. Simulation results verify the validity of these theoretically proven facets in terms of the efficacy and robustness of the proposed controller.
... Another research direction, multiphase homing theory also has its own problems. It is true that the multiphase homing theory is convenient to implement, for example, the homing method of X-38 aircraft recovery project [14]. Other researches such as the traditional multiphase homing method from Strahan [15], Yang [16] and Tao [17]. ...
... The multiphase homing strategy for X-38 PADS is presented in Ref. [15], as shown in Fig. 11. It is a typical method in the homing control and widely used in the airdrop experiment [14,15,17]. However, it can be clearly shown that the method could waste a lot of energy by tracking the energy management circle. ...
... It also has proved to be remarkably robust to uncertainty (Howitt, 2005;Smith and Berry, 2000). However, the achilles heel of NDI is the invertibility requirement on the plant's input distribution matrix G(x) (Ito et al., 2001). Even with the assumption that the inverse of this matrix exists for all states, x, it is often observed that G(x) becomes 'almost' singular for some x and hence can yield excessively large control signals. ...
... Even with the assumption that the inverse of this matrix exists for all states, x, it is often observed that G(x) becomes 'almost' singular for some x and hence can yield excessively large control signals. In fact, numerous examples exist in the literature of NDI controllers suffering significant losses in stability and performance as a result of actuator saturation (Ito et al., 2001) Actuator saturation has been studied for many years in the context of linear systems, and many anti-windup techniques have been proposed which limit the degradation of performance during saturation (see Hanus et al. (1987), Hippe and Wurmthaler (1999), Glattfelder and Schaufelberger (2003), Gomes da Silva Jr. et al. (2002) Crawshaw and Vinnicombe (2000) and for example). It is fair to say that major strides have been made in the treatment of this class of systems. ...
An anti-windup compensation method is proposed for a class of constrained nonlinear systems which, in the absence of saturation, are controlled by certain types of nonlinear dynamic inversion controllers. The anti-windup compensation scheme shares a similar architecture to that proposed by the authors in prior work for linear systems subject to saturation constraints. An encouraging aspect of the proposed scheme is that for globally exponentially stable systems, a particularly simple choice of anti-windup compensator exists and, moreover, this could be regarded as a "nonlinear" internal model control based antiwindup compensator. More generally, a framework for synthesising optimal antiwindup compensators is suggested, based on nonlinear partial differential matrix inequalities. Finally, a simple example illustrates the effectiveness of the scheme.
... Perhaps the most appealing aspect of NDI is that the design procedure inherently provides a nonlinear multivariable controller, thereby eliminating the necessity for further gain scheduling. The method has received significant attention from the aerospace community Reiner et al., 1996;Ito et al., 2001;Georgie, 2003;Snell et al., 1992;Snell, 1998;Escande, 1997;Bennani, 1998;G.Papageorgiou, 2001;C.Papageorgiou, 2005;Smith, 2000). The basis of NDI control is the feedback linearisation method which has become classical over the past 30 years (Isidori, 1995). ...
... However, the main impediment to the application of NDI is the invertibility requirement on the plant's input distribution matrix, G(x), which appears in the nonlinear affine system under consideration (Ito et al., 2001;Georgie, 2003). Even with the assumption that the inverse of this matrix exists for all states, x, it is often observed that G(x) becomes 'almost' singular for some x and hence excessively large control signals may occur. ...
A recently suggested general anti-windup (AW) compensation scheme is applied to the non-linear simulation model of the 1/16 scaled BAe Hawk aircraft model used for wind tunnel experiments. The Hawk is modelled as a nonlinear affine system subject to input constraints and has a primary control system consisting of an inner-loop nonlinear dynamic inversion controller and an outer-loop linear PID controller. To address the input constraints a recently introduced nonlinear L2 sub-optimal AW compensation method is applied and compared with a nonlinear version of the internal model control AW scheme. Nonlinear simulation results demonstrate the promise of the approach and indicate the superiority of the optimal AW scheme.
... Another widely used method is the linear robust controller [15,16]. However, a high-order robust controller is required, such as a fourteen-order controller to ensure the flying control system robustness for X-38 [17]. ...
This paper presents a L1 adaptive controller augmenting a dynamic inversion controller for UAV (unmanned aerial vehicle) carrier landing. A three axis and a power compensator NDI (nonlinear dynamic inversion) controller serves as the baseline controller for this architecture. The inner-loop command inputs are roll-rate, pitch-rate, yaw-rate, and thrust commands. The outer-loop command inputs come from the guidance law to correct the glide slope. However, imperfect model inversion and nonaccurate aerodynamic data may cause degradation of performance and may lead to the failure of the carrier landing. The L1 adaptive controller is designed as augmentation controller to account for matched and unmatched system uncertainties. The performance of the controller is examined through a Monte Carlo simulation which shows the effectiveness of the developed L1 adaptive control scheme based on nonlinear dynamic inversion.
... In [19], aircraft trajectory controller is designed for the aircrafts with model uncertainties and actuator faults by INDI. But for high performance demanded systems, large control signals usually cannot be avoided [20], which will cause the saturation of actuators. The control efficiency will be discounted and the stability of closed-loop system cannot be guaranteed in the presence of saturation. ...
In this paper, an anti-windup incremental nonlinear dynamic inversion (INDI) fault-tolerant scheme is proposed for flying wing aircraft with actuator faults, actuator saturation and uncertainties of aerodynamic parameters. An optimal anti-windup compensator based on nonlinear partial differential inequalities is used to compensate the actuator saturation. INDI is used to control the fault system and compensate the uncertainties of the flight dynamics. Control allocation strategy is designed in consideration of the control scheme and configuration of the control surfaces. The proposed control method can guarantee the bounded tracking of the reference signals. Simulation results are given to show the effectiveness of the proposed method.
... The controller design has been more and more difficult due to the complex nonlinearity, the strong coupling between control channels and the uncertainties of aerodynamic parameters in the non-power re-entry process. In existing methods such as robust control [2,3] and adaptive control [4,5] , the linearized model of the controlled system contains complex high-order derivative function, which is not convenient for practical application of engineering. Sliding mode control [6,7] compensates the influence of uncertainty by designing the virtual control quantity, but the sliding mode control is based on the upper bound of uncertainty and requires that its bound is known or the known function of the state variables, which is difficult to predict in practical applications. ...
For the hypersonic vehicle nonlinear attitude mode in reentry process with a strong coupling, aerodynamic parameter perturbations and non-deterministic, combine extended state observer and nonlinear law state error feedback, design the hypersonic vehicle MIMO-ESO ADRC attitude controller. Put interference such as uncertainty, coupling and parameter perturbations as “the sum of interference” ,use the extended state observer to estimate and dynamic feedback compensation, use nonlinear law state error feedback to inhibit residual of compensation. ADRC controller is charged without a precise model of vehicle , and without precise perturbation boundaries of aerodynamic parameters.Simulation results show that the MIMO-ESO ADRC attitude controller can overcome the impact of large-scale perturbations of interference and aerodynamic parameters, have good dynamic qualities and tracking capabilities, also have strong robustness.
... Meanwhile, the multiphase homing theory is more focused on the application. Like the X-38 aircraft recovery project [4]. The theories of the multiphase homing are similar [5]. ...
... According to [38,39], the attitude dynamic model of NSVs can be described bẏ ...
An adaptive fuzzy fault-tolerant tracking controller is developed for Near-Space Vehicles (NSVs) suffering from quickly varying uncertainties and actuator faults. For the purpose of estimating and compensating the mismatched external disturbances and modeling errors, a second-order sliding mode disturbance observer (SOSMDO) is constructed. By introducing the norm estimation approach, the negative effects of the quickly varying multiple matched disturbances can be handled. Meanwhile, a hierarchical fuzzy system (HFS) is employed to approximate and compensate the unknown nonlinearities. Several performance functions are introduced and the original system is transformed into one incorporating the desired performance criteria. Then, an adaptive fuzzy tracking control structure is established for the transformed system, and the predefined transient tracking performance can be guaranteed. The rigorous stability of the closed-loop system is proved by using the Lyapunov method. Finally, simulation results are presented to illustrate the effectiveness of the proposed control scheme.
... However, it is not easy to establish an accurate mathematical model of RLV and the controller design for RLV is challenging, as there are enormous amounts of model parameter uncertainties and external disturbances in which atmospheric disturbances is a significant component. [1][2][3] Several approaches for the controller design of RLV have been conducted over a period of time, for example, gain scheduling, 4 dynamic inversion technique, [5][6][7] back-stepping control, [8][9][10] theta-D, [11][12][13] statedependent Riccati equation strategy, 14,15 and trajectory linearization control. [16][17][18][19][20] However, these research works mentioned above have not taken atmospheric disturbances into account, which cannot be ignored both in theory and practice. ...
The controller design for reusable launch vehicles is challenging due to enormous amounts of model parameter uncertainties and atmospheric disturbances. This paper first derives six-degree-of-freedom model of a reusable launch vehicle with atmospheric disturbances. Next, four kinds of atmospheric disturbances are introduced and wind models are established respectively. For attitude control of the reusable launch vehicle, a nonsingular terminal sliding mode controller is designed with stability guaranteed. Finally, simulation results show a satisfactory performance for the attitude tracking of the reusable launch vehicle with atmospheric disturbances.
... In recent years, based on the single UAV model, many nonlinear control methods were proposed in carrier landing systems. In the field of control design, the latest research works Z. Zheng mainly focus on optimizing parameters [1][2][3], improving the control accuracy and robustness of the dynamic inversion control system [4], [5], improving the accuracy of sensor [6], reducing the noise of radar tracking and radio data link [7]. In addition, the fuzzy control method [8] was introduced into control system. ...
In this paper, relative motion model and control strategy for autonomous fixed-wing unmanned aerial vehicle(UAV) carrier landing are addressed. Firstly, a coupled six-degrees-offreedom (6-DOF) non-linear relative motion model is established from 6-DOF UAV and carrier models. Secondly, because of the under-actuated characteristic of two vehicles, the 6-DOF relative motion model is simplified to a four-degree-of-freedom (4-DOF) model to facilitate the control design. Thirdly, an adaptive sliding mode control law is proposed to track desired landing trajectory and maintain constant relative pitch and roll angles. Finally, simulation results demonstrate the effectiveness of the proposed control method.
... The DI control technique in turns allows to incorporate well known and established linear control methods. Moreover,to improve its robustness characteristics, several modifications were made to the basic DI control structure, see, e.g.,McFarland and Hoque (2000); Ito et al. (2001); Doman and Ngo (2002). Regardless of these attributes, DI has several shortcomings and limitations, including useful nonlinearity cancellation, large control effort, and computational issues that take place due to square matrix inversion. ...
This paper presents a generalized dynamic inversion attitude control design for satellite launch vehicle. Dynamic constraints that encapsulate the attitude tracking objectives are prescribed in the form of asymptotically stable differential equations in the attitude variables deviations from their reference trajectories. The constraint differential equations are evaluated along the vehicle's attitude trajectories and are inverted to solve for the control variables using the Greville formula for underdetermined algebraic systems. The associated null control vector is constructed using Lyapunov stability principle, and a delaying dynamic scaling factor is augmented in the involved Moore-Penrose generalized inverse for assured singularity avoidance and stable attitude tracking. The effectiveness of the proposed control scheme is shown by its evaluation on a 6DOF model of a satellite launch vehicle in the presence of various external disturbances.
... DI has been recently applied for many aircraft controllers. [11][12][13][14] Adams et al. designed a DI/µ-synthesis inner/outer loop control law for the thrust vectored F-18 in Refs. 15 and 16. ...
In this paper a robust control for aircraft longitudinal motion is presented. The Model-Error Control Synthesis (MECS) is developed for this application, which consists of a nominal controller with a model-error predictive filter. The control input is updated directly using the estimated model-error from a predictive filter to cancel the unmodeled dynamics or disturbance Inputs. The predictive filter is used only to estimate the model error, whereas an extended Kalman filter is used to provide state estimates from noisy measurements, The nominal control is chosen to be an LQR/dynamic inversion outer/inner loop controller for the longitudinal direction. The control law robustness to unmodeled dynamics and external disturbances is verified when MECS is active and not active. From various simulation results, MECS is shown to have better performance characteristics over the stand-alone nominal control. Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. rnr AIAA 2006-6050.
Channel modeling using the correlated double ring concept on Vehicular Network communications systems has been developed previously. However, no such model has been incorporated into the OFDM multi-carrier system to simulate the effect of the transmitter and receiver velocity on the received signal quality. User velocity on the transmitter and receiver side produces a Doppler effect that damages the signal orthogonality on OFDM. In addition, the speed of the transmitter and receiver also affect the Power Spectral Density of the received signal. This paper used the Correlated Double Ring channel modeling to simulate the transmitter and receiver movement against the Power Spectral Density parameter with non-moving scatterers. The IFFT output of the OFDM transmitter is consonant with the Channel Impulse Response (CIR) of the channel modeling output and coupled with the AWGN noise. Simulation is done by dividing the movement of sender and receiver into 3-speed regions, i.e. low speed, medium and high speed. Simulation results show the faster movement of the sender and receiver cause Doppler Shift larger and make the value of power spectral density parameters become getting muffled.
This paper presents the attitude control design of satellite launch vehicle based on the direct adaptive generalized dynamic inversion approach. The proposed adaptive generalized dynamic inversion approach encompasses the equivalent and the adaptive control elements. The equivalent (continuous) control part of adaptive generalized dynamic inversion is based on the conventional generalized dynamic inversion approach that comprises two noninterfering control actions, i.e. the particular part and the auxiliary part. In the particular part, dynamical constraint is prescribed in the form of time differential equation, which is evaluated along the vehicle attitude trajectories that encapsulates the control objectives and is inverted by utilizing Moore Penrose Generalized Inverse (MPGI). The singularity problem is solved by augmenting a dynamic scaling factor in the involved MPGI. In the auxiliary part, the null control vector is designed using the proportionality gain matrix, constructed by employing the Lyapunov function that guarantees global closed-loop asymptotic stability of the angular body rate dynamics. The adaptive (discontinuous) control part of adaptive generalized dynamic inversion is based on the sliding mode control with adaptive modulation gain, that provides robustness against tracking performance deterioration due to generalized scaling, system nonlinearities, and uncertainties, such that semi-global practically stable attitude tracking is guaranteed. External guidance loop based on the trajectory following method is designed, which reshapes the predefined pitch and yaw attitude profiles based on the respective normal and lateral positional errors, for acquiring the desired orbital parameters such as height, injection angle, orbital velocity, etc. To analyze the ascent flight trajectory, a detailed six-degrees-of-freedom simulator of a four-stage satellite launch vehicle is developed. The intensive numerical simulations are performed, which demonstrate the stability, robustness and the tracking capability of the proposed control and guidance methods in the presence of parametric uncertainties and external disturbances.
The nonlinear differential equations governing the motion of an aircraft are described in Chap. 3. For the plant analysis and control design, these equations are linearized around a certain operating point. Two sets of state variables appear to be clearly decoupled, each defining a specific mode of aircraft motion. The state variables involved in the longitudinal mode are the pitch rate q, the airspeed V
T
, the angle of attack α, and the pitch angle θ. The lateral-directional mode involves the state variables for the roll rate p, the yaw rate r, the sideslip angle β, and the roll angle φ. This chapter is dedicated to the analysis and control of the longitudinal motion of the aircraft and presents an architecture for the altitude controller, which uses robust NDI in all of the control loops. This chapter brings an innovative and practical approach for stability and robustness analyses of the plant undergoing the dynamic inversion process. Moreover, this chapter provides a systematic procedure for the selection of some uncertain model parameters involved in the controllers. Finally, a new nonlinear airspeed controller is also designed and presented.
An aircraft is intrinsically a nonlinear system. Therefore, if linear controllers are to be used in the aircraft flight control system, several linear controllers have to be designed and then gain-scheduled over the operating regime of the aircraft. However, recent nonlinear control techniques have made it possible to deal directly with the known nonlinearities of the aircraft dynamics, which yields a unique controller suitable for a wide range of operating conditions. This chapter describes the technique known as NDI and presents the architecture and the design procedure of the controllers used in the aircraft autopilot.
A novel robust adaptive control scheme is developed based on the characteristic modeling and discrete-time sliding mode control theory, for the attitude control of a hypersonic vehicle with parameter uncertainties and time-variation. The characteristic model of the hypersonic vehicle is derived considering the control objective and the vehicle feature firstly. Then a discrete-time sliding mode control law with a recursive form is proposed on the basis of the characteristic model, to enhance the control performance and system robustness to inevitable uncertainties. The main advantage of our proposed robust adaptive control scheme is that no knowledge of the boundaries and variation rate of parameter uncertainties is required in advance. Using the Lyapunov stability theory, the closed-loop system is proved stable and convergence properties of the system are assured. Finally, the simulations results demonstrate the effectiveness of the robust adaptive control scheme.
In this paper, the relative motion model and control strategy for autonomous unmanned aerial vehicle(UAV) carrier landing are addressed. Firstly, a coupled six degrees of freedom(6-DOF) non-linear relative motion model is established from (6-DOF) UAV and carrier models. Then the (6-DOF) relative motion model is simplified to the four degree of freedom (4-DOF) model to facilitate the control design, because of the underactuated characteristic of two vehicles. Secondly, the feedback linearization control law is proposed to control the UAV towards the aircraft carrier with constant forward, vertical velocity and lateral position and yaw angle. Finally, simulation results demonstrate the effectiveness of the proposed control system.
This paper addresses the 6 degrees-of-freedom control of a Mars Lander Module during its aerodynamic entry phase. It focuses on Robust Nonlinear Dynamics Inversion (RDI) technique. On one hand, RDI is shown to be efficient for translation controller synthesis. On the other hand, this technique is shown to be inappropriate for attitude control except for the gyroscopic torque cancellation which could be coupled with a robust static controller.Monte-Carlo simulations are performed to demonstrate the robust performance of the 6 DOF control software. The nonlinear simulator includes a realistic model of the atmosphere (European Mars Climate Database) and Mars gravity (J2).
It is obvious that the highly nonlinear nature of dynamic behavior in hypersonic vehicle system impose very challenging obstacles to the controller design. In this paper, we discuss the design of a novel type-2 fuzzy dynamic characteristic modeling method to attitude tracking control problem for hypersonic vehicles in gilding phase. The type-2 (T2) fuzzy logic is introduced into the characteristic modeling (CM) method. Unlike the traditional fuzzy dynamic modeling method normally with a fixed local linear model in every fuzzy subspace, our approach performs CM in subspace, which actually can handle the nonlinearity well while reducing the number of fuzzy rules. After dividing the whole restriction region into several subspaces, the whole nonlinear system can be regarded as a T2 fuzzy 'blending' of each individual characteristic model. Then this novel T2 fuzzy logic system modelled by decomposition of a complex nonlinear system into a collection of local CMs, can overcome the deficiencies of traditional fuzzy dynamic modeling and CM approaches. Therefore, it can achieve a better trade-off between tracking accuracy and convergence efficiency in the controller design for hypersonic vehicles. Simulation results under the conditions of certainty and uncertainty are given to show the effectiveness of the novel method for the attitude tracking control of hypersonic vehicles in gilding phase.
A novel adaptive neural network dynamic surface control is proposed for hypersonic vehicle. Based on the hypersonic vehicle model characteristics, the adaptive neural network dynamic surface control attitude controller and the neural networks velocity controller are designed, respectively. By combining Nussbaum gain function with decoupled backstepping, the problems of unknown control directions and control singularity from hypersonic vehicles model in strict-feedback form are simultaneously solved. Furthermore, the norm of the ideal weighting vector in neural network systems is considered as the estimation parameter, such that only one parameter is adjusted at each recursive step. Based on decoupled backstepping method and Lyapunov stability theorem, the semi-global stability of the flight control system is proved. Simulation results show that the new controller can guarantee the expected tracking performance. © 2015 Technical Committee on Control Theory, Chinese Association of Automation.
In this paper, a new sliding mode disturbance observer (SMDO) is developed using the terminal sliding mode technique. The SMDO is employed to estimate unknown external disturbances and modeling uncertainties in finite time. Based on the designed SMDO, a boundary layer adaptive sliding mode attitude control scheme is proposed for near-space vehicles (NSVs). The designed attitude control scheme can guarantee the satisfactory attitude tracking performance of the multi-input and multi-output (MIMO) attitude motion for the NSV subject to the time-varying disturbance. The rigorous stability of the closed-loop system is proved using the Lyapunov method. Finally, simulation results are presented to illustrate the effectiveness the proposed control scheme.
In this paper, to deal with the uncertain disturbances of near-space vehicles (NSVs), a new disturbance observer based on terminal sliding mode (TSMDO) technique and an adaptive backstepping sliding mode control scheme with the TSMDO are developed. The TSMDO can estimate the unknown external disturbances and the modeling uncertainties within the finite time and the differentiator is employed to avoid differential expansion problems in the backstepping control design. The rigorous stability of the closed-loop system is proved with Lyapunov method. Finally, simulation results illustrate that the proposed control scheme based on TSMDO can guarantee the satisfactory tracking performance of the multi-input and multi-output (MIMO) attitude motion for the NSV under the external unknown disturbance.
This paper presents nonlinear robust flight control strategies for the reentry vehicle which is nonlinear, coupling, and includes parameter uncertainties and external disturbances. Firstly, a finite-time second order sliding mode attitude control strategy is pointed out with the introduction of a nonsingular finite-time sliding mode manifold. By the proposed controller, the attitude tracking errors are mathematically proved to converge to zero within finite time and the chattering of sliding mode controller is alleviated without any deterioration of robustness and accuracy. For further alleviation of chattering, a smooth second order sliding mode attitude control strategy is then designed based on an improved nonsingular finite-time sliding mode manifold. Finally, the proposed strategies are applied to the attitude control of X-33 RLV in the reentry phase to verify the validity and robustness of the proposed strategies.
An attitude controller is designed for a reusable launch vehicle (RLV) during the re-entry phase with input constraint, model uncertainty and external disturbance. A control-oriented model with matched and unmatched uncertainty is first derived for the control design. To proceed, the control scheme is designed taking advantage of the robust stability property of sliding mode control, the compensation ability of the non-linear disturbance observer (DOB) technique and the systematic design procedure of the backstepping technique. An additional system, the states of which are used for controller design and stability analysis, is constructed to handle the input constraint. The time derivative of the virtual control input is considered as an uncertainty to eliminate the ‘explosion of terms’ problem inherent in backstepping control. Moreover, the proposed method alleviates the chattering problem of the traditional sliding mode backstepping control scheme. Stability analysis of the composite control scheme consisting of the designed controller and the DOB is performed via Lyapunov theory. Finally, simulation results are compared to show that the proposed method is able to achieve better tracking performance and tackles input constraints more effectively than an adaptive filter backstepping control scheme.
The work developed and presented in this thesis focuses on the design, validation and comparison of different nonlinear control solutions allowing an airship to navigate autonomously. To accomplish this task, a six-degrees-of freedom nonlinear model of the airship is developed based on the Lagrangian equations, reproducing the airship response to actuator and wind disturbances inputs. The linearization of this model for trim conditions over the flight envelope results in the known decoupling of the longitudinal and lateral motions, and allows a thorough analysis of the airship control design problem over the entire aerodynamic range. The conditions are then set to propose alternative nonlinear control solutions so as to have a single control law valid for different missions, independent of the flight region, and robust to realistic wind disturbances. The control methodologies developed in this work for the airship autonomous flight are Gain Scheduling, Dynamic Inversion and Backstepping. Besides the analysis of specific problems inherent to the design and implementation of each controller, desired performance criteria are also defined, allowing the comparison of the different solutions. This assessment, based on simulation results for complete flight missions defined from take-off to landing, and considering realistic wind disturbances, is important in order to establish the viability of the controllers implementation onboard the experimental airship platform. This thesis is part of the research made in the area of nonlinear flight control of airships for the AURORA and DIVA projects of the Institute of Mechanical Engineering (IDMEC) in Instituto Superior Tecnico, Technical University of Lisbon.
Purpose
– The purpose of this paper is to present a sliding mode attitude controller for reusable launch vehicle (RLV) which is nonlinear, coupling, and includes uncertain parameters and external disturbances.
Design/methodology/approach
– A smooth second-order nonsingular terminal sliding mode (NTSM) controller is proposed for RLV in reentry phase. First, a NTSM manifold is proposed for finite-time convergence. Then a smooth second sliding mode controller is designed to establish the sliding mode. An observer is utilized to estimate the lumped disturbance and the estimation result is used for feedforward compensation in the controller.
Findings
– It is mathematically proved that the proposed sliding mode technique makes the attitude tracking errors converge to zero in finite time and the convergence time is estimated. Simulations are made for RLV through the assumption that aerodynamic parameters and atmospheric density are perturbed. Simulation results demonstrate that the proposed control strategy is effective, leading to promising performance and robustness.
Originality/value
– By the proposed controller, the second-order sliding mode is established. The attitude tracking error converges to zero in a finite time. Meanwhile, the chattering is alleviated and a smooth control input is obtained.
The paper presents an attitude control problem of reusable launch vehicles in reentry phase. The controller is designed based on synthesizing robust adaptive control into backstepping control procedure in the presence of input constraint, model uncertainty, and external disturbance. In view of the coupling between the states of translational motion and the states of attitude motion, the control-oriented model is developed, where the uncertainties do not satisfy linear parameterization assumption. The time derivative of the virtual control input is viewed as a part of uncertain term to facilitate the analytic computations and avoid the ‘explosion of terms’ problem. The robust adaptive backstepping control scheme is first proposed to overcome the uncertainty and external disturbance. The robust adaptive law is employed to estimate the unknown bound of the uncertain term. Furthermore, the attitude control problem subjects to input constraint is studied, and the constrained robust adaptive backstepping control strategy is proposed. Within the Lyapunov theory framework, the stability analysis of the closed-loop system is carried out, and the tracking error converges to a random neighborhood around origin. Six-degree-of-freedom reusable launch vehicle simulation results are presented to show the effectiveness of the proposed control strategy. Copyright © 2015 John Wiley & Sons, Ltd.
An integrated guidance and control scheme is developed for next generation of reusable launch vehicle (RLV) with the aim to improve the flexibility, safety and autonomy. Firstly, an outer-loop optimal feedback reentry guidance law with online trajectory reshaping capability is designed. Then, a novel reentry attitude control strategy is proposed based on multivariables smooth second-order sliding mode controller and disturbance observer. The proposed control scheme is able to guarantee that the guidance commands generated from the guidance system can be tracked in finite time. Furthermore, a control allocation is integrated in the system in order to transform the control moments to control surface deflection. Finally, some representative simulation tests are conducted to demonstrate the effectiveness of the proposed integrated guidance and control strategy for six-degree-of-freedom RLV.
This paper addresses the 6 degrees-of-freedom control of a Mars Lander Module during its aerodynamic entry phase. It focuses on Robust Nonlinear Dynamics Inversion (RDI) technique. On one hand, RDI is shown to be efficient for translation controller synthesis. On the other hand, this technique is shown to be inappropriate for attitude control except for the gyroscopic torque cancellation which could be coupled with a robust static controller. Monte-Carlo simulations are performed to demonstrate the robust performance of the 6 DOF control software. The nonlinear simulator includes a realistic model of the atmosphere (European Mars Climate Database) and Mars gravity (J2).
A robust nonlinear controller is designed for a complete Unmanned Aerial Vehicle UAV mission employing a combination of dynamic inversion and H ∞, control. An inner-loop/outer-loop control structure is employed to synthesize a full envelope nonlinear controller for a complete UAV mission. In the inner loop, a modified time-scale separation approach is employed to feedback linearize the nonlinear model. In the outer loop, an H ∞ control procedure is used to achieve performance and robustness design goals. The reference flight envelope is a complete mission. The uncertainties considered include uncertainty in UAV model parameters such as the stability derivatives, weight, center of gravity location, and airspeed variations. Atmospheric disturbances and measurement noise are also included. Since robustness and disturbance rejection must be verified, flight simulation under a variety of uncertainties is performed using MATLAB/Simulink. Copyright © 2006 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
Using a model-following neuro-adaptive approach, a nonlinear controller has been designed in this paper for a practical reusable launch vehicle that assures robust tracking of guidance commands despite having uncertainties in the plant model. The overall control design is carried out in two steps. First a nominal design is carried out based on the philosophy of dynamic inversion (for RCS) and optimal dynamic inversion (for aerodynamic control), which assures tracking of the guidance commands for the nominal plant with an optimal control allocation strategy. In this design phase, the bank angle and angle-of-attack commands as issued by the guidance algorithm are first converted to equivalent roll and pitch commands respectively, while simultaneously assuring turn-coordination through the necessary yaw rate command generation. These body rate commands are then tracked in an inner-loop by generating the necessary control surface deflections. Next, an adaptive control is designed that enforces the inner-loop body rates of the actual plant to track the closedloop body rates of the nominal plant. This overall control design structure retains the simplicity in design while simultaneously assuring robust performance of the controller. The control design has been carried out using the full Six-DOF model of the vehicle in velocity frame that imbeds the spherical and rotating earth effects in the vehicle dynamics (which is in harmony with the dynamics used for the vehicle guidance). The promising simulation results with the Six-DOF dynamics along with realistic constraints like actuator dynamics, control bounds, RCS constraints etc. clearly demonstrate the good command following as well as robustness of the overall design approach presented in this paper, making it a viable technique to be implemented in a real vehicle.
Dynamic Inversion is a design technique used to synthesize flight controllers whereby the set of existing dynamics are cancelled out and replaced by a designer selected set of desired dynamics. The output of such an inner loop controller is the control input required to achieve the desired response. The desired dynamics essentially form a loop-shaping compensator that affects the closed-loop response of the entire system. This paper attempts to quantify the particular form of desired dynamics which produce the best closed-loop performance and robustness in a Dynamic Inversion flight controller. Four candidate forms of desired dynamics which invert the short period dynamics are evaluated. These four include a proportional, proportional integral, flying qualities, and a ride qualities form of desired dynamics. Longitudinal controllers are synthesized for the prototype X-38 Crew Return Vehicle using a linear model at a selected point in the flight envelope. Pole placement is used to synthesize a robust outer loop around the dynamic inversion inner loop in order to provide closed-loop stability. The resulting closedloop performance is evaluated in the time domain, and in terms of frequency dependent singular values, quadratic cost and a passenger ride comfort index. Of the desired dynamics presented here, results indicate that the ride quality compensation dynamics provide the best overall system performance in terms of both time domain and frequency domain responses.
Among the more than 20 coupling derivatives, two of them are the most important ones in representing the major effects of lateral-directional motion on longitudinal motion. These are the sideslip coupling derivative (X) and the bank angle coupling derivative (Z). Other derivatives are negligible when compared to these two derivatives. In this paper, these two derivatives are fully derived and discussed. Using these coupling derivatives, the state space representation of the 6 DOF coupled motion is presented. Then, eight new transfer functions giving longitudinal state variables (forward velocity, angle of attack, pitch angle, and pitch rate) with respect to aileron and rudder deflections are derived. As a sample application, the coupled state space formulation is used to design the optimal control system of an airplane to do a complete 3-D mission.
Dynamic inversion (DI) is a control synthesis technique that steers the system states to track the desired trajectory by cancelling the original dynamics. However, it is impossible to achieve perfect cancellation of the original dynamics in real applications, robustness problem must be considered in designing DI controller. In this paper, robust dynamic inversion (RDI) control strategy based on sliding mode control (SMC) is proposed. Different from the previous works of achieving robustness in DI controller, the proposed RDI controller is designed inherently robust. Moreover, robust observer strategy based on the RDI controller is also proposed in this paper. Numerical simulations with application to the AV-8A Harrier aircraft demonstrate the effectiveness of the proposed method. © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
The dynamical modeling of an Unmanned Air Vehicle (UAV) forms the heart of its simulation. The equations of motion can take five different forms: 1) nonlinear fully coupled, 2) nonlinear semi-coupled, 3) nonlinear decoupled, 4) linear coupled, and 5) linear decoupled. In the fully coupled equations of motion, two new coupling stability derivatives ( ! D C , ! L C ) are incorporated. The main purpose of this paper is to compare each version of the equations of motion, and to demonstrate positive and negative features of each form. This paper also compares the five different dynamical models using eight distinct UAV missions. It analyzes them from three different aspects, and gives recommendations for which model is best for a specified mission. MATLAB/Simulink is used to implementation the simulation. The final results are compared by analyzing the resulting trajectories and control deflections.
An inner-loop/outer-loop dynamic inversion control structure with output redefinition in the inner loop is employed to synthesize a full envelope nonlinear autopilot for an UAV. In the inner loop, the time-scale separation approach was employed to feedback linearize the nonlinear model. In the outer loop an H∝ loop shaping procedure, incorporating characteristics of classic loop-shaping and H∝ design, is used to achieve performance and robustness design goals. Since the controller structure must be capable of attenuating disturbances while handling uncertainties in the dynamic model, a two degree-of-freedom controller is used. The reference flight envelope is a 3-D trajectory including take-off, climb, cruise, coordinated turns, descent, approach, and landing. The uncertainties considered include wind gusts, measurement noise, atmospheric altitude and airspeed variations, and uncertainty in UAV model parameters such as the stability derivatives and weight. Since the method must then be verified, its flight simulation is performed using MATLAB/Simulink. Copyright © 2005 by the American Institute of Aeronautics and Astronautics. Inc. All rights reserved.
Model Predictive Control (MPC) has the advantage of including constraints into the optimization. By combining MPC with Feedback Linearization (FBL) it is possible to use linear, discrete time MPC algorithms with nonlinear models. This way the constraints on the control systems, thrusters and aerodynamic surface deflections, and on the attitude of the vehicle can be integrated into the controller synthesis. The main disadvantage of this type of controller is the lack of robustness when uncertainties enter the system. This paper therefore improves the nominal FBL-MPC controller by replacing the Quadratic Programming algorithm with a min-max MPC technique involving Linear Matrix Inequalities. Uncertainties on the nonlinear aerodynamic coefficients are used as a source of disturbances. By mapping these uncertainties to a linear state space model as used by the MPC controller, a polytopic set of uncertainty models is created that can be used by the min-max algorithm to minimize the performance cost over the worst case model from the uncertainty set. This approach will be illustrated using a model of the X-38 re-entry vehicle flying along a predefined trajectory with active input and state constraints. The results show that the robustness is improved significantly when compared to the nominal controller with the same uncertainties. Copyright © 2005 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
The similarities and differences of dynamic inversion control and model-following control laws are examined. For the forms of these control laws assumed in this paper it is shown that dynamic inversion may be considered a special case of model-following. For any given dynamic inversion control law there is a model-following control law that achieves exactly the same response and therefore is in every way equivalent to it. This same model-following control law may be modified in its error dynamics without changing the desired response implied by the dynamic inversion law. The modification in error dynamics may be used to improve the tracking of the desired response in the presence of modeling errors. (Author)
Dynamic inversion is presented as a general tool for designing decoupling control laws for multivariable systems, with the difference that instead of using conventional, static, full-state feedback, a dynamic compensator is used. This has the advantages that fewer measurements are required and that the designer has better control over the bandwidth of inner feedback loops. This feature is important because it allows the control law to be made less sensitive to the effects of unmodeled sensor and actuator dynamics, which can be destabilizing. The main focus is on applications to linear systems, although some results may extend to nonlinear systems. The issue of zero dynamics and nonminimum phase is discussed in the linear context. Three examples are presented relating to decoupling of lateral-directional dynamics for highly maneuverable aircraft.
Minimax methods are proposed for the analysis and design of controllers for the best controls with the worst initial conditions, worst parameter changes with specified quadratic norms, and worst disturbances with specified integral-square norms. The worst initial conditions are the only forcing functions; disturbances are regarded as an added set of feedback controls whose magnitudes are limited by negative weights in the performance index. The minimax value of the performance index is easily calculated as the maximum eigenvalue of a steady-state Lyapunov or Riccati matrix. There is a lower bound on the disturbance weights in the performance index; at this bound, the controller design is identical to the H ∞ controller design. There is also upper bound on the norm of the parameter changes; at this value, the closed-loop system goes unstable, and the corresponding parameter change vector is almost the same as the corresponding vector obtained by “real μ” analysis — the only difference being the use of a quadratic norm instead of an infinity norm.
The objective of this paper is to propose both feedforward and feedback controllers designed by the Linear Quadratic Gaussian and Loop Transfer Recovery (LQG/LTR) methodology. The adoption of a two-degree-of-freedom controller with feedforward and state-feedback controllers can improve the disadvantages of high gain for only feedback controllers designed by the LQG/LTR method. This paper also derives the algorithms which can take all the time-domain, frequency-domain, and robust design techniques into a unified method. An example of F-16 lateral autopilot design is given, which shows that the proposed method is more robust to measurement noise and more practical in applications, and has better steady-state responses in time domain.
This paper presents a methodology for the design of flight controllers for aircraft operating over large ranges of angle of attack, The methodology is a combination of dynamic inversion and structured singular value (mu) synthesis. An inner-loop controller, designed by dynamic inversion, is used to linearize the aircraft dynamics. This inner-loop controller lacks guaranteed robustness to uncertainties in the system model and the measurements; therefore, a robust, linear outer-loop controller is designed using mu synthesis. This controller minimizes the weighted H-infinity norm of the error between the aircraft response and the specified handling quality model while maximizing robustness to model uncertainties and sensor noise. The methodology is applied to the design of a pitch rate command system for longitudinal control of a high-performance aircraft. Nonlinear simulations demonstrate that the controller satisfies handling quality requirements, provides good tracking of pilot inputs, and exhibits excellent robustness over a wide range of angles of attack and Mach number, The linear controller requires no scheduling with night conditions.
In this paper we overview the so-called Linear-Quadratic-Gaussian method with Loop-Tranfer-Recovery (LQG/LTR). Our objective is to provide a pragmatic exposition, with special emphasis on the step-by-step characteristics for designing multivariable feedback control systems.
Techniques for the design of control systems for manually controlled, high-performance aircraft must provide the following: (1) multi-input, multi-output (MIMO) solutions, (2) acceptable handling qualities including no tendencies for pilot-induced oscillations, (3) a tractable approach for compensator design, (4) performance and stability robustness in the presence of significant plant uncertainty, and (5) performance and stability robustness in the presence actuator saturation (particularly rate saturation). A design technique built upon Quantitative Feedback Theory is offered as a candidate methodology which can provide flight control systems meeting these requirements, and do so over a considerable part of the flight envelope. An example utilizing a simplified model of a supermaneuverable fighter aircraft demonstrates the proposed design methodology.
Dynamic-inversion-based flight control laws present an attractive alternative to conventional gain-scheduled designs for high angle-of-attack maneuvering, where nonlinearities dominate the dynamics. Dynamic inversion is easily applied to the aircraft dynamics requiring a knowledge of the nonlinear equations of motion alone, rather than an extensive set of linearizations. However, the robustness properties of the dynamic inversion are questionable especially when considering the uncertainties involved with the aerodynamic database during post-stall flight. This paper presents a simple analysis and some preliminary results of simulations with a perturbed database. It is shown that incorporating integrators into the control loops helps to improve the performance in the presence of these perturbations.
Dynamic inversion is a technique for control law design in which feedback is used to simultaneously cancel system dynamics and achieve desired dynamic response characteristics. However, dynamic inversion control laws lack robustness to modeling errors if improperly designed. This paper examines a simple linear example, control of roll rate about the body axis of high performance aircraft, to illustrate some robustness problems which may occur with a simple dynamic inversion control law. The paper demonstrates how structured singular value synthesis techniques can be used to enhance the robustness properties of the dynamic inversion controller.
Actuator rate saturation is an important factor adversely affecting the stability and performance of aircraft flight control systems. It has been identified as a catalyst in pilot-induced oscillations, some of which have been catastrophic. A simple design technique is described that utilizes software rate limiters to improve the performance of control systems operating in the presence of actuator rate saturation. As described, the technique requires control effectors to be ganged such that any effector is driven by only a single compensated error signal. Using an analysis of the steady-state behavior of the system, requirements are placed upon the type of the loop transmissions and compensators in the proposed technique. Application of the technique to the design of a multi-input/multi-output, lateral-directional control system for a simple model of a high-performance fighter is demonstrated as are the stability and performance improvements that can accrue with the technique.
Translational Motion Control of VSTOL Aircraft Using Nonlinear Dynamic Inversion
- P R Smith