Browse Topic: Attitude control
Modern space exploration relies more and more on vehicles that carry expansive antenna arrays, a trend driven by steady progress in the field of space engineering. The study centers on constructing a coupled model that accounts for structural rigidity, flexibility, and thermal effects in a sizable orbital unfurling parabolic antenna, along with designing its orientation regulation system. A blended coordinate formulation and a heat-based discretization approach are employed to derive the motion and temperature-governing equations of the assembly. A unified pointing control method is then developed by merging classical feedback control principles with torque compensation control. Computational experiments involving a spacecraft fitted with such a dish antenna are executed. The findings reveal that sudden heat fluctuations, exemplified by the transition out of Earth’s umbra, trigger oscillatory behavior in the antenna framework and disrupt the satellite’s angular position. The designed orientation regulator successfully counteracts these thermally driven perturbations, even when the system representation contains inaccuracies.
In response of the stringent requirements of high-resolution optical remote sensing satellites for attitude control stability and micro-vibration environment, this study focuses on the friction dynamics modeling and simulation analysis of large-angular-momentum flywheels across a wide speed range. To quantify the impact of flywheel friction disturbances on satellite attitude pointing accuracy, friction torque characteristic data were first acquired through ground testing over the full speed range, and a high-fidelity static fitting mathematical model was constructed using the least squares method. To further elucidate the nonlinear friction mechanisms in the low-speed regime—particularly during speed zero-crossings—the LuGre dynamic friction model was introduced, and its core parameters were identified and optimized. Comparative results demonstrate that the model exhibits high-precision characterization capabilities, with a deviation of less than 0.0015 Nm from measured data within the critical low-speed range of ±50 rpm. Subsequently, the established friction model was integrated into a satellite attitude control closed-loop system for digital simulation to quantitatively evaluate the disturbance of friction torque on system steady-state accuracy. The results indicate that under the closed-loop control compensation mechanism, the impact of flywheel friction on satellite attitude pointing accuracy is effectively suppressed, with the attitude angle deviation maintained within 0.002°, thereby meeting the engineering requirements for high-precision remote sensing imaging. This research provides a theoretical basis and data support for the optimization design of flywheel systems and the refined design of satellite attitude controllers.
To solve a problem that ignition anomaly can’t be detected in time, based on the thermal equilibrium equation, the space heat flow, heater heating, propellant combustion, and thermal radiation to cryogenic space are considered to build an accurate ignition temperature method for the 10 N thruster by using on-orbit true temperature. Further, considering the error of measuring the thermistor, an envelope model for the 10 N thruster ignition temperature is established. Based on the above, a detection method for the 10 N thruster ignition anomaly of on-orbit satellites is proposed. The accuracy of the method is relatively high, and the absolute error is less than 3 degrees Celsius. An anomaly can be quickly detected when the 10N thruster ignition temperature deviates from the normal trend by 3–5 degrees celsius. The method is applied to a DFH-3 satellite, and the maximum difference of 10 N thruster ignition temperature between the theoretical values calculated by the proposed method and the measured values is only 2.72 degrees celsius. It has been proven that the prediction accuracy of the proposed method is high. It plays an important role in discovering the 10N thruster ignition anomaly in time and ensuring the success of satellite orbit or attitude control.
This paper presents an initial handling qualities analysis of an Electric Vertical Take-Off and Landing (eVTOL) hexacopter. The analysis uses the Distributed Electric Propulsion Simulation (DEPSim), developed by Penn State University (PSU) and the Comprehensive Hierarchical Aeromechanics Rotorcraft Model (CHARM), developed by Continuum Dynamics, Inc. (CDI). The study focuses on evaluating a generic AAM hexacopter performing Handling Qualities Task Elements (HQTE) as defined by the DOT / FAA. A trajectory controller was developed to enable simulation of prescribed flight paths, allowing automated simulation of four HQTEs: Heliport Approach, Hovering Turn and Hold, Pirouette, Lateral Reposition and Hold. Design modifications incorporating lateral mast tilt and Direct Side Force Control (DSFC) were implemented to enhance yaw control and ride qualities. Piloted simulations were conducted at the PSU rotorcraft flight simulation facility using DEPSim, employing an Attitude Command Attitude Hold (ACAH) architecture with mode switching to Translational Rate Command / Position Hold (TRC / PH) and TRC plus DSFC modes. Two of the four HQTEs were tested in piloted simulations. Though formal ratings were not collected at this time, pilot commands and performance indicated that TRC / PH and TRC plus DSFC modes enhance handling qualities over ACAH mode. The DSFC control law was found to have substantially reduced roll attitude, which could potentially enhance visual cueing, pilot comfort, and pilot-perceived handling qualities.
The transition phase of eVTOL aircraft poses a challenge in balancing energy efficiency and stability. This study presents the development and evaluation of an automatic flight control system for eVTOL transition phases, focusing on minimizing energy consumption while ensuring robust performance. The control architecture implements a hybrid response type combining Translational Rate Command below 5 knots and Acceleration Command Speed Hold above 5 knots, with control allocation dynamically adjusted based on airspeed and rotor shaft angle. Stability analysis reveals surge mode instability at high shaft angles due to negative speed stability derivatives, stabilized through carefully tuned feedback control. The system demonstrates Level 1 handling qualities against bandwidth, quickness, and disturbance rejection criteria when evaluated against MIL-DTL-32742 and MIL-STD-1797B standards. Simulation results verify the control system's ability to maintain precise acceleration/deceleration rates and attitude control while ensuring passenger comfort through limited pitch excursions. The control strategy achieves minimum energy transitions by locking rotor shaft angles to optimal schedules while avoiding excessive hub moments. Flight test maneuvers developed specifically for conversion phases confirm the system's capability to execute efficient transitions within defined performance boundaries. This research establishes a framework for certifiable eVTOL flight control systems that balance energy efficiency with robust performance across diverse flight regimes.
This study investigates the fault tolerance of a large-scale coaxial quadrotor Electric Vertical Takeoff and Landing (eVTOL) under motor failure through high-fidelity software-in-the-loop (SIL) simulations using PX4-Gazebo environment. The objective is to evaluate the vehicle's ability to maintain flight stability and complete critical missions under various propulsion failure scenarios, without the control system being explicitly aware of which motors have failed. Four motor failure cases-single, two adjacent, two diagonally opposite, and three distributed motor failures-were introduced during takeoff, hover, cruise, and hover under crosswind missions. Results show that the eVTOL maintained controllability and mission completion under all scenarios, with increasing levels of performance degradation under more severe failures. Notably, considerable yaw instabilities of about 10 degrees occurred under two diagonally opposite motor failures. The highest thrust demands after motor failures were observed during cruise mission, with some motors demanding about 80% to 90% of their maximum throttle. Hover under crosswind revealed compounded challenges in attitude control during descent under severe failure cases compared to calm weather. These findings underscore the robustness of the integrated control system and vehicle configuration in managing motor failure scenarios.
The flight dynamics of quadrotor UAVs exhibit strong nonlinearity and inter-axis coupling, presenting significant challenges to the design of their flight control systems. This paper proposes a new attitude controller integrates feedback linearization model together with predictive control (FLMPC) for the quadrotor UAVs. Model predictive control (MPC) can determine the optimal control input of the system under constraints. Feedback linearization transforms the nonlinear flight dynamics model of the quadrotor UAV into a linear decoupled system, simplifying the solution of the optimal control problem. To validate the effectiveness of the controller, a model of the controlled object and control law is established for attitude control simulation, and comparisons are made with linear MPC and adaptive MPC. The results show that the proposed flight control method can achieve attitude control of the quadrotor UAV with a low computational burden.
The presence of a slung-load during the flight of a quadrotor generates swing effects that can greatly influence the dynamics of the quadrotor. These effects have the potential to threaten the stability of the system’s attitude. This study presents a disturbance compensation strategy that is designed based on the utilization of an adaptive harmonic extended state observer (AHESO) in order to solve this problem and achieve precise attitude control. To derive the aforementioned algorithm, a comprehensive mathematical model for the quadrotor-slung-load system is built. The periodic features of disturbance are derived by considering the movement of the slung-load. Subsequently, by taking the periodic features of the disturbances into account, the AHESO for accurate disturbance estimation is designed. In this observer, an online frequency estimator for the harmonic disturbances is introduced. Lyapunov theory is introduced to examine the stability of the AHESO. In addition, backstepping-based approach is employed to design an attitude controller for accurate attitude tracking of the quadrotor. Analysis are conducted regarding of the entire stability of the quadrotor-slung-load system. In the end, the effectiveness of the designed method is verified by comparative simulations.
The quad tilt rotor unmanned aerial vehicle (QTR-UAV) is a nonlinear tough coupled system with more redundant actuators. Its flight dynamics are affected by internal and external disturbances such as rotor tilt motion and wind disturbance. It is difficult to establish an accurate flight dynamics model and design a stable flight control system. In this paper, the flight dynamics model affected by internal and external disturbances is established based on the flight principle. The unknown disturbances are estimated through the adaptive neural network extended state observer. The redundant control strategy design is used to overcome the influence of rotor tilting. The rotor tilting process is stably controlled in the transition corridor as the flight speed changes. The channel control law algorithm uses a nonlinear state error feedback regulator to compensate for the influence of internal and external disturbances. The channel control outputs are distributed according to the manipulation strategy. The QTR-UAV flight safety and stability are ensured. A flight simulation system for the QTR is constructed, which is implemented with an adaptive radial basis function neural network active disturbance rejection control (RBF-ADRC). The attitude control simulation is conducted and some simulation results are presented. It is demonstrated that the designed RBF-ADRC flight control system can be effectively and reliably applied to the flight control of QTR, and its tracking accuracy, anti-interference and robustness are better than traditional active disturbance rejection controller and PID controller.
Aiming at the problem of flight control of unmanned helicopter with sensor faults, a linear active disturbance rejection (LADRC) fault-tolerant control method based on RBF neural network is proposed. The integral variable of the state output signal is introduced to augment the original system. On this basis, the RBF neural network is used to estimate the unknown sensor fault, and the LADRC technology is used to obtain the system state quantity and the estimated value of the total internal and external disturbances, and the system is transformed into a standard integral series structure without disturbance. Based on Lyapunov stability theory, it is proved that the closed-loop system is globally asymptotically stable. Finally, the attitude control of unmanned helicopter with different types of sensor faults is simulated. The results show that the RBF-LADRC fault-tolerant control method proposed in this paper can effectively overcome the influence of sensor faults, and has strong anti-disturbance ability and good robustness.
This paper describes the design, development, and tethered flight testing of a quad-cyclocopter weighing 25 kilograms and 1.8 m x 1.4 m x 0.8 m in dimension. This cyclocopter features four cantilevered cyclorotors with a unique five-bar blade pitching mechanism. The cyclorotor design is chosen through systematic parametric studies using an in-house 2-D computational fluid dynamics (CFD) solver. Based on the parametric studies, the final design selected is a 6-bladed cyclorotor with 0.67 chord-to-radius ratio, symmetric NACA 0015 airfoil, and pitch amplitude of 45-degree because it provided high thrust and power loading (thrust/power) at a low operating rotational speed of 700 RPM. The cyclorotor blades are manufactured with a foam core and carbon fiber skin resulting in lightweight blades with large bending and torsional stiffness. The rotor supporting structure and transmission is designed to be lightweight and resilient to large centrifugal loads and dynamic torques, respectively. A minimum mass airframe is designed and constructed from carbon fiber tubes to handle the vibratory loads from cyclorotors, mount electronics and batteries, and provide attachment points for the carbon fiber cowling. The attitude control strategy utilizes a combination of RPM control and thrust vectoring of the four cyclorotors (eight independent control parameters), resulting in an overactuated system. A custom autopilot uses a closed-loop cascaded proportional-integral-derivative (PID) controller to stabilize the aircraft in hovering flight. The vehicle has been flight tested while tethered and achieved lift off.
The operation of helicopters on ships is one to most challenging tasks due to adverse weather conditions, the lack of visible cues, turbulent airwakes behind the ship and a moving confined landing spot on the ship. Currently, only a very limited number of pilot assistance systems are available to ease helicopter ship deck landings. The focus of this paper is the evaluation of a Head-Down Display (HDD), a Head-Mounted Display (HMD) and two different Attitude Command Attitude Hold (ACAH) flight control architectures for ship deck landings based on piloted simulation. A ship deck landing scenario at the research flight simulation facility Air Vehicle Simulator (AVES) has been extended to include turbulent ship airwakes from high-fidelity Computational Fluid Dynamics (CFD). The pilot assistance systems have been implemented at the simulator and evaluated by four helicopter pilots. In particular, the results show a favorable potential of the Head-Mounted Display and the flight control architectures.
In the field of space technology, Centre Suisse d’Electronique et de Microtechnique (CSEM) has been a partner of the European Space Agency (ESA) for many years. One focus of their joint research is eliminating vibration emissions originating from components aboard satellites. In addition to limiting the precision of attitude control for satellites, these micro-vibrations lead to higher energy consumption and (in the case of imaging missions) cause deterioration of image quality.
To achieve Level 1 Handling Qualities, Aeronautical Design Standard ADS-33E-PRF requires an Attitude Command Attitude Hold or Translational Rate Command response-type in Degraded Visual Environments while allowing a rate response in Good Visual Environments. The authors describe the design and analysis of a Blended Command Model that may offer the precision of the former and the aggressiveness of the latter. The command model, comprised of a single flexible transfer function and its parameter scheduling functions, produces attitude, rate, and blended responsetypes as functions of cyclic stick position. The authors explain the command model’s design considerations, test considerations, and its behavior through time domain and frequency domain perspectives. This effort precedes a handling qualities assessment aboard the U.S. Army’s JUH-60A RASCAL aircraft.
In application, the Aeronautical Design Standard for the handling qualities of military rotorcraft, ADS-33E-PRF, provides the means to effectively predict rotorcraft handling qualities via validated criteria and demonstrate actual handling qualities in flight test using mission task elements. Besides a definition, a note that rotorcraft shall have no tendencies, and a note regarding Attitude Command Response-Types and gain bandwidth frequency, the topic of pilotinduced oscillations (PIO) is not addressed via specific criteria or flight test techniques. As the use of full authority fly-by-wire flight control continues to expand in Vertical Takeoff and Landing (VTOL) aircraft, the likelihood of encountering PIO will also expand. In the fixed wing world where PIO has been commonplace, at least in developmental test if not operations, predictive analytical methods that can also be used for detection of PIO in realtime have been developed, which can also be applied to rotorcraft. Furthermore, recent time-frequency domain methods that have been developed to differentiate VTOL piloting techniques are directly applicable to PIO scenarios. Using a flight test database generated with the UH-60L in-flight simulators at the U.S. Naval Test Pilot School (USNTPS), this paper explores the utility of these methods for identifying rotorcraft PIO tendencies.
When firing artillery, there is typically a maximum angle that the platform cannot exceed relative to the Earth plane. This is due to the large recoil forces involved and the risk of destabilizing the platform the weapon is mounted to. Mobile systems are particularly sensitive to this as the attitude of the platform relative to Earth is constantly changing. A simple solution is to add pitch and roll sensors directly to the platform. However, many mobile systems already have an assortment of sensors that can be used to calculate the platform attitude.
NASA’s Small Spacecraft Technology Program is on the countdown clock to advance communications and proximity maneuvering capabilities for CubeSats with the Integrated Solar Array and Reflectarray Antenna (ISARA) mission.
The complexity of helicopter flight dynamics combined to an increasing demand to improve its mission performance are two sources of difficulties for the efficient design of helicopter flight control systems. Handling qualities requirements, as specified in ADS-33 standards (Ref. 1) provide a useful guideline for control laws design. The right selection of the criteria in relation with the helicopter planned missions (MTEs) and the usable cues to realize them (UCE) can ensure an optimized design of the control system. However, designing robust control laws with such constrained requirements remains a challenging task. Recent activities at ONERA have led to the development of an approach to integrate handling qualities requirements in robust control design techniques. As a first step Handling Qualities charts are generated using an equivalent model of a conventional PID controller for an Attitude Command Attitude Hold (ACAH) control law. After what the charts are used to feed robust control design tools. The paper presents the charts, their validation process and their use in ONERA robust control design environment.
In designing of the Attitude Control System (ACS) is important take into account the influence of the structure’s flexibility, since they can interact with the satellite rigid motion, mainly, during translational and/or rotational maneuver, damaging the ACS pointing accuracy. In the linearization and reduction of the rigid-flexible satellite mathematic model, usually one loses some important information associated with the satellite true dynamical behavior. One way to recovery this information is include to the ACS design parametric and not parametric uncertainties of the system. The H infinity control method is able to take into account the parametric uncertainty in the control law design, so the controller becomes more robust. This paper presents the design of a robust controller using the H infinity control technique to control the attitude of a rigid-flexible satellite. The satellite model is represented by a flexible beam connected to a central rigid hub considering a set of parametric uncertainties. The results of the simulations have shown that the control law designed with the H-infinity control method was robust enough as to support the action of uncertainty and control the rigid flexible satellite attitude and suppressing vibrations.
The present contribution aims at providing a comprehensive illustration of a structured approach to the design, implementation and testing of a new rotor state measurement system for rotorcraft applications. This effort has been carried out in the framework of a Clean Sky collaborative project in which the novel sensor system plays a fundamental role by enabling the real-time estimation of non-measurable quantities that govern the rotorcraft running acoustic emission, in view of external noise alleviation. Furthermore, the availability of the new sensor system capable to accurately capture rotor blade motion allows the derivation of enhanced attitude control laws based on rotor state feedback. We detail the complete process that led to the full-scale development of a stereoscopic vision-based measurement system mounted on the rotor head, which has been fully integrated on board a prototype helicopter for ground and flight testing.
The vast majority of the U.S. Army's helicopter fleet consists of aircraft initially developed in the 1960s and 1970s and which were designed based on the handling qualities and flight control requirements of the time for flight in good visual environments (GVE). The Army today uses helicopters at night and in brownout and other degraded visual environment (DVE) conditions but with the same control laws of the original models; the major exception being the CH-47F and MH-47G DAFCS, which have been highlighted as a successful partial authority flight control system upgrade to provide improved handling qualities. The U.S. Army Aviation Development Directorate–AFDD has partnered with the U.S. Army Utility Helicopter Program Office's Futures Team and the RDECOM DVE Mitigation Program to further develop and test the UH-60 Modernized Control Laws (MCLAWS). Previous work implemented a model following control system architecture which provided an attitude command/attitude hold response-type for hover and low speed flight. This system demonstrated improved handling qualities as compared to the UH-60L SAS/FPS rate command response-type. This paper documents work to integrate an outer-loop position hold with velocity command mode into the MCLAWS. Flight testing of the MCLAWS with position hold demonstrated Level 1 Cooper-Harper handling qualities ratings in simulated DVE conditions. Finally, landing logic has been integrated into the MCLAWS to support DVE landing flight testing.
A joint research project (2010-2014) between Delft University of Technology and Boeing Mesa was conducted in SIMONA Research Simulator (SRS) at Delft University with the goal to develop advanced flight control laws for handling qualities (HQs) improvements of the Apache AH-64 helicopter. The goal of the present paper is to concentrate on implementation and simulator testing of modern control laws for Apache's AH-64D Longbow helicopter to provide improved handling qualities for hover and low speed flight in degraded visual environment. The paper will implement an "Incremental Nonlinear Dynamic Inversion (INDI)" controller into the Boeing's FlyRT AH-64 Apache baseline model for the existing partial authority stability augmentation system (SAS). The INDI will be used to provide both attitude command attitude hold and translational rate command response types based on the requirements in ADS-33E. Implementation of the INDI into Apache's FlyRT proved to be challenging because the model was not affine in the control inputs. Actuator dynamics and rotor dynamics (most likely flapping dynamics) caused the designed INDI controller to overcontrol the helicopter. The paper proposes three adaptations of the INDI methodology that prove able to stabilize the helicopter and can be used as an alternative to a full authority, fly-bywire, control system upgrade. Handling qualities evaluations conducted in SIMONA simulator for the ADS-33 hover and pirouette maneuvers demonstrated that the INDI resulted in improving from Level 2 to Level 1 HQs, reducing the pilot workload when compared to the Apache's legacy flight control system.
This paper describes the design, development and flight testing of a meso-scale cyclocopter. Weighing only 29 grams, the present vehicle is the smallest cycloidal rotor based aircraft ever built. Unlike the previous cyclocopters, the current prototype utilizes a novel, light weight (3 grams) cycloidal rotor design, with cantilevered blades, having semi-elliptical planform shape and no exposed rotor shaft. To minimize bending deflections the blades use a unique, lightweight (0.15 grams each) but high strength-to-weight ratio unidirectional carbon-fiber based structural design and are fabricated using a specialized manufacturing process. The cycloidal rotor design was chosen through systematic performance measurements conducted using a custom-built miniature three-component force balance. Based on experimental parametric studies, a 4-bladed rotor and symmetric blade kinematics with pitch amplitude of 45° provided the highest thrust and power loading (thrust/power) and was used in the final rotor design. The airframe is fabricated using a combination of carbon-fiber and state-of-the-art 3D printing techniques. The attitude control strategy utilizes a combination of rpm-control of the two cycloidal-rotors/tail-rotor and thrust vectoring of the cycloidal rotors. The control strategy is implemented on a custom-built 1.3 gram autopilot, which uses a closed-loop proportional-derivative controller for hover stability. The vehicle has been systematically flight tested by tuning the feedback gains and has demonstrated stable hovering flight.
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