Nominal Flight Conditions Research Articles

Intersatellite and Downlink Agile Attitude Maneuvers

This research is focused on the problem of agile attitude maneuvering, aimed at the precise pointing of a satellite forming a typical constellation in low Earth orbit. We consider two different operational scenarios: (a) pointing toward a specific ground station, located on the Earth surface (for downlink data routing), and (b) pointing toward a companion satellite (for establishing an intersatellite connection). The two preceding operational requirements can both be formulated as attitude tracking problems. In this study, we use an inertia-free nonlinear attitude control algorithm based on rotation matrices and possessing remarkable stability properties, in conjunction with a pyramidal array of single-gimbal control momentum gyroscopes. Numerical simulations, in both nominal and nonnominal flight conditions, demonstrate that the attitude control architecture proposed in this work is effective for the purpose of performing agile attitude maneuvering, aimed at precise pointing during downlink and intersat data routing.

Sizing and Analysis of a Lift-Plus-Cruise Aircraft with Electrified Propulsion

This paper demonstrates an integrated approach to the sizing and analysis of all-electric, hybrid-electric, and turboelectric variants of a lift-plus-cruise urban air mobility vertical takeoff and landing concept using the Parametric Energy-based Aircraft Configuration Evaluator. In lieu of assuming a point mass performance model, the aircraft is trimmed explicitly for nominal and off-nominal flight conditions for sizing propulsion system components and evaluating mission energy usage. In lieu of using historical tail volume ratio guidelines for stabilizer sizing, an explicit dynamic stability analysis is incorporated to determine the stabilizer size that achieves specified bare airframe handling qualities targets. The effects of varying battery specific energy, C-rate, wing loading, hover disk loading, mission range, and cruise speed on the sizing of the three propulsion architectures are explored. A comparison of the weight breakdown of the three propulsion variants when sized to the same maximum takeoff mass is presented. The sensitivity of the sizing results to the requirement of full vs reduced payload capability at high, hot hover conditions is also assessed. In addition to being insightful, the presented results reinforce the need to incorporate more physics-based analyses, additional stability and control considerations, and configuration optimization studies, which form natural avenues for follow-on work.

Waypoint Navigation of Quadrotor using Deep Reinforcement Learning

This paper proposes a Reinforcement Learning (RL) based technique to develop a simple neural network controller for the task of waypoint navigation in quadrotors. In this paper, the application of Twin Delayed Deep Deterministic (TD3) Policy Gradient algorithm for high and low-level control implementation for quadrotors is discussed. The proposed methods are tested on high fidelity Gym-Pybullet-Drones simulator. The effectiveness of the methods developed is validated through numerical simulations. The simulation results indicate that both control policies are successful in navigating through the assigned waypoint, with the low-level controller being accurate in the nominal flight conditions. In the presence of disturbance inputs, the high-level controller performs better when compared to the low-level controller.

Neighboring Optimal Guidance and Constrained Attitude Control for Accurate Orbit Injection

Accurate orbit injection represents a crucial issue in several mission scenarios, e.g., for spacecraft orbiting the Earth or for payload release from the upper stage of an ascent vehicle. This work considers a new guidance and control architecture based on the combined use of (i) the variable-time-domain neighboring optimal guidance technique (VTD-NOG), and (ii) the constrained proportional-derivative (CPD) algorithm for attitude control. More specifically, VTD-NOG & CPD is applied to two distinct injection maneuvers: (a) Hohmann-like finite-thrust transfer from a low Earth orbit to a geostationary orbit, and (b) orbit injection of the upper stage of a launch vehicle. Nonnominal flight conditions are modeled by assuming errors on the initial position, velocity, attitude, and attitude rate, as well as actuation deviations. Extensive Monte Carlo campaigns prove effectiveness and accuracy of the guidance and control methodology at hand, in the presence of realistic deviations from nominal flight conditions.

Enhancing VTOL Multirotor Performance With a Passive Rotor Tilting Mechanism

This article discusses the benefits of introducing a simple passive mechanism to enable rotor tilting in Vertical Take-Off and Landing (VTOL) multirotor vehicles. Such a system is evaluated in relevant Urban Air Mobility (UAM) passenger transport scenarios such as hovering in wind conditions and overcoming rotor failures. While conventional parallel axis multirotors are underactuated systems, the proposed mechanism makes the vehicle fully actuated in SE(3), which implies independent cabin position and orientation control. An accurate vehicle simulator with realistic parameters is presented to compare in simulation the proposed architecture with a conventional underactuated VTOL vehicle that shares the same physical properties. In order to make fair comparisons, controllers are obtained solving an optimization problem in which the cost function of both systems is chosen to be equivalent. In particular, the control laws are Linear-Quadratic Regulators (LQR), which are derived by linearizing the systems around hover. It is shown through extensive simulation that the introduction of a passive rotor tilting mechanism based on universal joints improves performance metrics such as vehicle stability, power consumption, passenger comfort and position tracking precision in nominal flight conditions and it does not compromise vehicle safety in rotor failure situations.

Optimizing sizing schedules of aircrew helmets using machine learning techniques

Introduction: Aircrew equipment assemblies are life critical equipment worn by the aircrew to ensure protection from various life-threatening environmental conditions that may occur in nominal or off-nominal flight conditions. It is important for the sizing system of body worn aircrew equipment to be divided into different size classes such that it is as representative of existing groups in the population as feasible. Machine learning techniques provide a powerful tool to develop efficient sizing systems that are representative of actual body types existing in the population. The objective of the study was to use machine learning clustering techniques to identify naturally existing body types to formulate an optimal sizing system for an aircrew helmet. Material and Methods: The standard sizing schedule of an aircrew helmet using head length and head breadth was studied. An iterative semi-supervised machine learning protocol called K-means clustering was used to identify naturally occurring clusters of head types within the population. The cluster boundaries were identified to develop the final sizing schedule. Analysis of variance (ANOVA) with post hoc analysis was carried out on the four head lengths and head breadths. Results: Four clusters of head type were identified using K-means clustering. This led to a sizing schedule which can be descriptively called “Short-Narrow,” “Short-Broad,” “Long-Narrow,” “Long-Broad.” ANOVA showed a statistically significant difference between the four clusters for both head length and head breadth. Conclusion: Consideration of several variables in sizing schedules often fails to meet desired fitment. Standard sizing methods use linear partitions on individual control parameters. In reality, the proportional variation between parameters is not linear. Machine learning tools help in identification of naturally occurring clusters within the database considering several variables at the same time. This study uses machine learning techniques to identify existing groups in a population and develop a final sizing schedule. Similar analysis for other aircrew equipment assemblies can be carried out to determine sizing schedules which assist in design and fitment.

Design and Evaluation of UAV Flow Angle Estimation Filters

This paper presents the design, implementation, and evaluation of four filters for the estimation of angle of attack (AOA) and angle of sideslip (AOS) of small unmanned aerial vehicles (UAVs). Specifically, two novel filters (a complementary filter and an extended Kalman filter) are proposed and evaluated without using direct flow angle and Global Positioning System measurements; two existing AOA/AOS filters are also implemented and evaluated. All filters are designed with minimal inputs and states to ensure the ease of implementation, simplicity of tuning, and computation efficiency. Both simulation and UAV flight test results show the performance of the proposed filters. Especially, flight test results from two different UAVs (a T-tail UAV and a flying wing UAV) show that the root mean square errors of estimated inertial AOA and AOS are less than 1.5 $^\circ$ under nominal flight conditions and around 2 $^\circ$ under aggressive maneuvers compared with direct flow angle measurements.

Modeling, Trim Analysis, and Trajectory Control of a Micro-Quadrotor with Wings

This paper presents the modeling, control design, and efficiency analysis of a micro-quadrotor aerial vehicle with airfoils. We derive the equations of motion for a micro-quadrotor (length 0.15 m and mass 0.03 kg) outfitted with two symmetric airfoils that generate lift during forward transit. A trim analysis is presented to determine nominal flight conditions. Analysis of the model facilitates design of a controller that stabilizes the modified quadrotor through transitions from hover to forward fixed-wing flight. The performance of the design and control approach is analyzed through a series of typical flight profile simulations. The controller is able to track a velocity trajectory in the presence of aerodynamic uncertainties. The simulations are also used to determine the efficiency of the aircraft compared to a standard quadrotor. Results suggest that, during forward flight, the airfoils result in up to a 35% energy savings depending on the aspect ratio and forward speed of the vehicle. Added airfoil weight, however, causes a 45% energy loss during hover. These results indicate that the neutral benefit speed is in the range 3–5 m/s depending on the aspect ratio of the airfoil, which is suggestive of mission flight profiles that best utilize the added benefit of airfoils on a micro-quadrotor.

Benchmark Aerostructural Models for the Study of Transonic Aircraft Wings

Since its introduction, the NASA Common Research Model has proved a useful aerodynamic benchmark for predicting computational-fluid-dynamics-based drag and aerodynamic design optimization. The model was originally conceived as a purely aerodynamic benchmark, and as such the wing geometry corresponds to the deflected shape at its nominal flight condition. However, interest has been growing to extend this model to aeroelastic studies. Unfortunately, because of its predefined deflection, the model is not suitable for aeroelastic analysis and design. To address this issue, an undeflected Common Research Model is defined through an inverse design procedure that includes the outer mold line geometry of the undeflected wing and the corresponding internal wingbox structure. Additionally, because modern transport aircraft are trending toward higher-aspect-ratio wing designs, a higher-aspect-ratio variant of the Common Research Model wing is developed to assess next-generation wing designs. This variant has an aspect ratio of 13.5 and is designed by using buffet-constrained multipoint aerostructural optimization. The purpose of these models is to provide publicly available benchmarks for aeroelastic wing analysis and design optimization.

Integrated Navigation, Guidance and Control System and Validation

The navigation, guidance and control (NGC) system of Reusable Launch Vehicle – Technology Demonstrator (RLV-TD) is the most complex system ever flown in ISRO’s launches. It includes a navigation system employing traditional inertial navigation system along with global positioning system aiding and radar altimeter, ceramic servo accelerometer package, angle of attack-based control, different control systems, including secondary injection thrust vector control, reaction control system and aero control surfaces. Closed loop simulations of integrated NGC system of RLV-TD are of utmost importance to ensure mission success by assessment of system performance in nominal and off-nominal flight conditions. Validation is achieved through progressive evaluations in different levels of integrated simulation, namely autonomous simulations, software in loop simulation, on-board processor in loop simulation, hardware in loop simulation and actuator in loop simulation using Iron Bird test facility. This article discusses the overall NGC system, challenges faced during the realization of the simulation test bed, including system design to validate the new on-board elements, commissioning of Iron Bird facility and validation through various phases of simulation. The outcome of these simulations played a crucial role in bringing about improvements and helping in the evolution of the on-board system to the final configuration, ensuring the success of the mission.

Accelerated Flight Control LAW Clearance Using Stores Grouping Concept

Current fighter aircraft are demanded to feature multirole capability with large number of different external stores to be integrated on the aircraft. These stores represent air to air, air to surface weapons etc., which have wide range of aerodynamic characteristics as well as different mass properties (CG, mass, inertia). This conflicts with the requirement of adequate stability, optimum handling qualities, superior agility and robustness characteristics over the entire flight envelope. Especially with an unstable aircraft it is impossible to fulfill these requirements for all stores with just one set of control laws. Therefore, automatic reconfiguration in control laws based on similar flight dynamic characteristics is carried out. Stores grouping concept is often used to reconfigure the flight control laws with store information in order to normalize/optimize the flight dynamics characteristics and corresponding support towards flight clearance. The flight clearance is a tedious procedure as a host of criteria need to be evaluated at a large number of nominal flight conditions for different airframe configurations, multiples thereof with tolerances and failures. This paper describes how various stores configurations can be grouped together in order to accelerate the clearance process.

Autonomous Flight Envelope Estimation for Loss-of-Control Prevention

A nonlinear aircraft dynamics modeling and system identification approach is formulated and used to derive a rapid, convex optimization-based algorithm for the estimation of aerodynamic force coefficients from noisy sensor data, with uncertainty quantification. Furthermore, with the aerodynamic force coefficients identified, a computationally fast method for generating the corresponding aircraft trim envelope given safe limits on achievable thrust and angle of attack is presented. This unified approach achieves the online Bayesian inference of the aircraft aerodynamic performance capability while running in the background as the aircraft is controlled independently by the pilot or automation system. The resulting information is readily incorporated into the determination of the extended safe maneuvering envelope, pilot displays, and automation algorithms for flight planning, trajectory generation, and guidance: all to help maintain safe aircraft operations under both nominal and off-nominal flight conditions.

Variable-Time-Domain Neighboring Optimal Guidance, Part 2: Application to Lunar Descent and Soft Landing

In recent years, several countries have shown an increasing interest toward both manned and automatic lunar missions. The development of a safe and reliable guidance algorithm for lunar landing and soft touchdown represents a very relevant issue for establishing a real connection between the Earth and the Moon surface. This paper applies a new, general-purpose neighboring optimal guidance algorithm, proposed in a companion paper and capable of driving a dynamical system along a specified nominal, optimal path, to lunar descent and soft landing. This new closed-loop guidance, termed variable-time-domain neighboring optimal guidance, avoids the usual numerical difficulties related to the occurrence of singularities for the gain matrices, and is exempt from the main drawbacks of similar algorithms proposed in the past. For lunar descent, the nominal trajectory is represented by the minimum-time path departing from the periselenium of a given elliptic orbit and arriving at the Moon with no residual velocity. Perturbations arising from the imperfect knowledge of the propulsive parameters and from errors in the initial conditions are considered. At specified, equally spaced times the state displacements from the nominal flight conditions are evaluated, and the guidance algorithm yields the necessary control corrections. Extensive robustness and Monte Carlo tests are performed, and definitely prove the effectiveness, robustness, and accuracy of the new guidance scheme at hand, also in comparison with the well-established linear tangent steering law.

Simple and efficient numerical tools for the analysis of parachutes

Purpose – The purpose of this paper is to describe a set of simple yet effective, numerical method for the design and evaluation of parachute-payload system. The developments include a coupled fluid-structural solver for unsteady simulations of ram-air type parachutes. The main features of the computational tools are described and several numerical examples are provided to illustrate the performance and capabilities of the technique. Design/methodology/approach – For an efficient solution of the aerodynamic problem, an unsteady panel method has been chosen exploiting the fact that large areas of separated flow are not expected under nominal flight conditions of ram-air parachutes. A dynamic explicit finite element solver is used for the structure. This approach yields a robust solution even when highly nonlinear effects due to large displacements and material response are present. The numerical results show considerable accuracy and robustness. Findings – A simple and effective numerical tool for the analysis of parachutes has been developed. Originality/value – An analysis code has been developed which addresses the needs of ram-air parachute designers. The software delivers reasonably accurate results in a short time using modest hardware. It can therefore assist the design process, which nowadays relies on empirical methods.

A Novel Design of Landing Gear Oleo Strut Damper Using MR Fluid for Aircraft and UAV’s

Most landing gears used in aircraft employ very efficient oleo-pneumatic dampers to absorb and dissipate the impact kinetic energy of the aircraft body frame. A single-acting shock absorber is most commonly used in the oleo strut that has a metering pin extending through the orifice, which can vary the orifice area upon compression and extension of the strut. This variation is adjusted by shaping the metering pin so that the strut load is fairly constant under dynamic loading. In this paper, it is proposed to further change the damping coefficient as a function of time in order to achieve a semi-active control of the aircraft vibrations during landing by using Magnetorheological (MR) fluid in the Oleo. With the metering pin designed for a nominal flight condition, further variation in the fluid viscosity would help achieve the optimal performance in off-nominal flight conditions. A simulation approach is employed to show the effect of different profiles for viscosity variation in the MR fluid. The utility of such a damper can be very well exploited to include different criteria such as the landing distance after touchdown. This type of system can be used also in Unmanned Aerial Vehicle (UAV) application where the focus of design may be to accomplish the task without the consideration of passenger comfort.

Complementary Information Distribution to Improve Team Performance in Military Helicopter Operations: An Experimental Study

This study compared a complementary information distribution, that is, purposeful distribution to individual crewmembers of the information needed for their individual tasks, to the current distribution where all crewmembers can access all information in the flight deck. The complementary information distribution is intended to foster a complementary mental model that promotes each team member's access to information according to his or her immediate tasks. The hypothesis—that the complementary information distribution would improve performance and team interaction—was tested in a high-fidelity military helicopter flight simulator with 20 Army aviators. The complementary information distribution was found to have no effect or a positive effect (and no negative effects) on measures of individual's task performance and error rates, with greater positive effects on most measures in off-nominal compared to nominal flight conditions. In addition, measures of team communication and awareness of fellow crewmembers' information needs and workload also found only no effect or positive effects with a complementary information distribution.

Hybrid adaptive ascent flight control for a flexible launch vehicle

For the purpose of improving trajectory tracking performance of a launch vehicle under adverse flight conditions, this article describes the development of a hybrid adaptive ascent flight control architecture. The hybrid adaptive control approach augments linear feedback signals with contributions from both direct and indirect adaptive control elements. In addition, signal filters are incorporated into the feedback loop to prevent harmful interaction between the flight control system and structural bending modes. The performance of the hybrid adaptive flight controller is compared to that of a typical gain-scheduled linear feedback controller in a high-fidelity Ares I ascent simulator obtained from the NASA Marshall Space Flight Center. Results from simulations which consider nominal flight conditions show that the hybrid adaptive controller is able to maintain a tracking error magnitude that is comparable to or lower than its classical proportional–integral–derivative (PID) counterpart for the vast majority of the ascent phase. However, when a 10 per cent thrust reduction is applied to the first-stage rocket motor, the PID controller experiences significantly elevated error that diverges rapidly at the end of the simulation. In contrast, the hybrid adaptive controller steadily maintains a tracking error magnitude that is frequently 50 per cent or less of the corresponding PID value during the latter stages of ascent.

Attitude Controller Design and Test of Korea Space Launch Vehicle-I Upper Stage

This paper introduces the upper stage attitude control system of KSLV-I, which is the first space launch vehicle in Korea. The KSLV-I upper stage attitude control system consists of two electro-hydraulic actuators and a reaction control system using cold nitrogen gas. A proportional, derivative, and integral controller is designed for the electro-hydraulic thrust vectoring system, and Schmidt trigger ON/OFF controllers are designed for the reaction control system. Each attitude controller is designed to have enough stability margins. The stability and performance of KSLV-I upper stage attitude control system is verified via hardware in the loop tests. Hardware in the loop tests are accomplished for perturbed flight conditions as well as nominal flight condition. The test results show that the attitude control loop of KSLV-I upper stage is very stable and the attitude controllers perform well for all flight conditions. Attitude controllers designed in this paper have been successfully applied to the first flight of KSLV-I on August 25, 2009. The flight test results show that all attitude controllers of the KSLV-I upper stage performed well and satisfied the accuracy specifications even during abnormal flight conditions.

A simulation environment for design and testing of aircraft adaptive fault‐tolerant control systems

PurposeThe purpose of this paper is to present the development of a Matlab/Simulink‐based simulation environment for the design and testing of indirect and direct adaptive flight control laws with fault tolerant capabilities to deal with the occurrence of actuator and sensor failures.Design/methodology/approachThe simulation environment features a modular architecture and a detailed graphical user interface for simulation scenario set‐up. Indirect adaptive flight control laws are implemented based on an optimal control design and frequency domain‐based online parameter estimation. Direct adaptive flight control laws consist of non‐linear dynamic inversion performed at a reference nominal flight condition augmented with artificial neural networks (NNs) to compensate for inversion errors and abnormal flight conditions following the occurrence of actuator or sensor failures. Failure detection, identification, and accommodation schemes relying on neural estimators are developed and implemented.FindingsThe simulation environment provides a valuable platform for the evaluation and validation of fault‐tolerant flight control laws.Research limitations/implicationsThe modularity of the simulation package allows rapid reconfiguration of control laws, aircraft model, and detection schemes. This flexibility allows the investigation of various design issues such as: the selection of control laws architecture (including the type of the neural augmentation), the tuning of NN parameters, the selection of parameter identification techniques, the effects of anti‐control saturation techniques, the selection and the tuning of the control allocation scheme, as well as the selection and tuning of the failure detection and identification schemes.Originality/valueThe novelty of this research efforts resides in the development and the integration of a comprehensive simulation environment allowing a very detailed validation of a number of control laws for the purpose of verifying the performance of actuator and sensor failure detection, identification, and accommodation schemes.

GENETIC ALGORITHM FOR COUPLED RLV TRAJECTORY AND GUIDANCE OPTIMIZATION

This paper deals with an innovative application of Genetic Algorithm tools in order to perform a global optimization of the hypersonic re-entry trajectory of a reusable launch vehicle, or RLV, with Guidance scheme and Control performance model in-the-loop. Driving idea is to rely at termination of GA process on a reference trajectory in total adequation with Guidance and Control settings and performances and compliant with all mission and GNC requirements for both nominal and off-nominal flight conditions. Illustration of developed principles is made considering a medium L/D entry vehicle.