Flight Envelope Research Articles

A Comparison of Isolated and Ducted Fixed-Pitch Propellers under Non-Axial Inflow Conditions

A strong interest in highly-efficient, small-scale propeller configurations can be recognized, especially due to the currently growing number of and usage possibilities for unmanned aerial vehicles (UAVs). Although a variety of different propulsion concepts already exist on the market or are discussed in the literature, there is still a demand for a systematic investigation to compare such configurations, in particular, small-scale propellers with a fixed pitch, which are analyzed in this work. Therefore, different configurations of small-scale propellers with a fixed pitch are analyzed in this paper. They were operated as isolated single propellers and as ducted propellers in a cylindrical wing. Furthermore, due to their flight envelope, UAVs are likely to operate at highly inclined inflow conditions and even under reverse inflow. These non-axial inflow conditions have a major influence on the flow field around a propeller. In order to investigate this influence, all analyses were performed at a range of inflow angles in relation to the propeller axis from αdisc=0° to 180°.

Reliability-based Design Optimization of Classical Wing Aeroelasticity

Flutter speed of aircraft is very important and needs to be firstly specified before a certification applied for a new aircraft by airworthiness regulator to make sure that the aircraft is free from flutter in its flight envelope. By assuming geometrical and physical parameters known, the speed is usually estimated from deterministic analyses in a design stage. In practice, some parameters are finitely measured by observing, especially for the geometrical parameters, material properties and so on due to the random in nature, which causes uncertainty of information often called uncertainties. The purpose of this paper is to combine reliability analysis and optimum design of aeroelastic aircraft wing. The classical two-dimensional wing with a typical airfoil section is used as an example in this study. To quantify uncertainty in the design of flutter speed, the discrete-time aero-elastic model and worst-case scenario are applied. Furthermore, the comparison between optimum design with/without reliability is provided in this study. The results show the proposed technique leads to the flutter speed being more conservative and realizable compared with the traditional technique.

Numerical investigation of compressibility effect on dynamic stall

A rotorcraft main-rotor blade experiences a broad range of the Mach number in high-speed forward flights or rapid maneuvers. Near the boundary of the flight envelope, the rotor blade often encounters severe dynamic stall which limits the overall aerodynamic performance. The effects of the flow compressibility on dynamic stall due to the variation of the freestream Mach number are studied here. A rotor airfoil, VR-12, is computationally investigated for both static and dynamic stall conditions in the relevant range of the Mach number, 0.2 to 0.4. For the small enough Mach number such as 0.2, both static and dynamic stalls are significantly influenced by flow separation due to adverse pressure gradient with little compressibility effect. For the highest Mach number 0.4, compressibility near the leading edge is no longer negligible, forming a local supersonic pocket there. The shock-induced separation occurs near the leading edge, which dominates both the static and dynamic stall at the high Mach number. At the intermediate Mach number 0.3, it is observed that both adverse pressure gradient and weak shock affect the stall. Unsteady compressible Reynolds-averaged Navier-Stokes equations are numerically solved with the Spalart-Allmaras turbulence model in the current study. Numerical results are compared to relevant wind-tunnel test data.

Interplay of unsteady aerodynamics and flight dynamics of transport aircraft in icing conditions

Airframe icing causes significant degradation of aerodynamic characteristics and influences the flight safety. Wind tunnel study of longitudinal steady and unsteady aerodynamic characteristics of a transport aircraft in icing conditions is carried out in order to develop mathematical model of aerodynamics in the extended flight envelope. The wind tunnel results are validated through flight tests conducted for the real aircraft. Large, glaze-horn ice shapes, corresponding to holding flight phase, are considered. Influence of an ice protection system as well as its failure is examined. Effect of icing on the unsteady aerodynamics characteristics is studied not only through wind tunnel tests but also via analysis of subsequent influence on the flight dynamics of the aircraft. The conducted study shows that the ice shapes of the holding phase lead to reduced stall angle of attack (AoA), maximum lift, and longitudinal damping. Flight dynamics analysis demonstrates that dangerous aircraft behaviour in the form of high AoA departure and limit cycle oscillations (LCO) can be observed at smaller elevator deflections for the iced aircraft. The account for icing influence on the unsteady aerodynamics in the flight dynamics simulations revealed degradation of the dynamic response and deterioration of phase portraits of the system. Even for small AoA and elevator deflection, the aircraft might be trapped into the basin of attraction of high-AoA LCO. In addition, incorporating icing effects in unsteady aerodynamics manifest larger amplitude of LCO.

Continuous morphing trailing-edge wing concept based on multi-stable nanomaterial

Morphing technology is one of the most effective methods to improve the flight efficiency of aircraft. Traditional control surfaces based morphing method is mature and widely used on current civil and military aircraft, but insufficiently effective for the entire flight envelope. Recent research on morphing wing still faces the challenge that the skin material for morphing should be both deformable and stiff. In this study, a continuous morphing trailing-edge wing with a new multi-stable nano skin material fabricated using surface mechanical attrition treatment technology was proposed and designed. Computational fluid dynamics simulation was used to study the aerodynamic performance of the continuous morphing trailing-edge wing. Results show that the lift coefficient increases with the increase of deflection angle and so does the lift-drag ratio at a small angle of attack. More importantly, compared with the wing using flaps, the continuous morphing trailing-edge wing can reduce drag during the morphing process and its overall aerodynamic performance is improved at a large angle of attack range. Flow field analysis reveals that the continuous morphing method can delay flow separation in some situations.

Improved Model for On-Board Real-Time by Constructing Empirical Model via GMM Clustering Method

The method of constructing an empirical model is used to compensate the deviation between the output of the on-board real-time model and the engine measurement parameters, and improve the parameter tracking and estimation performance of the on-board adaptive model in the full flight envelope. Due to the large amount of data acquired online, the clustering method based on Gaussian mixture model is implemented to realize data compression for offline training and updating the empirical model. The present empirical model is applied to the on-board adaptive model of civil large bypass ratio turbofan engine. The simulation results show that the empirical model based on Gaussian mixture model can reduce the output error of on-board real-time model, and the accuracy of the health parameter estimation and engine component fault isolation performance of the on-board real-time adaptive model with empirical model are improved.

A bio-inspired flight control strategy for a tail-sitter unmanned aerial vehicle

Wingbeat behavior and intermittent flight path are the two main characteristics of many birds. In this paper, to improve the efficiency of energy use and cruise range, a bio-inspired intermittent flight strategy with a whole flight envelope has applied to a tail-sitter aircraft. A total energy control system based transition control law has been proposed. The energy efficiency is investigated in terms of energy consumption per unit distance of different cruising modes, and the effectiveness and stability of proposed flight mode transition control law are verified by simulation. The mean mechanical power in flap-gliding flight is reduced compared with steady flight.

Robust proportional incremental nonlinear dynamic inversion control of a flying-wing tailsitter

Tailsitter unmanned aerial vehicles are utilized extensively nowadays since they merge advantages of both fixed-wing unmanned aerial vehicles and rotary-wing unmanned aerial vehicles. However, their attitude control suffers from unknown nonlinearities and disturbances due to the wide flight envelope. To solve the problems, a robust attitude controller based on a newly designed flying-wing tailsitter is proposed in this paper. By employing the angular acceleration feedback to compensate unmodeled dynamics, the proportional incremental nonlinear dynamic inversion control law is first developed. The proportional incremental nonlinear dynamic inversion strengthens the conventional nominal gain incremental nonlinear dynamic inversion with a proportional term to reflect the change of the angular acceleration more directly. Therefore, the tailsitter has a quicker response and performs better in suppressing model uncertainties and external disturbances. Since the angular acceleration is difficult to measure in practice, an angular acceleration estimation method is then established to provide accurate signals for the proportional incremental nonlinear dynamic inversion. The signals are derived as complementary results of model prediction method and direct differential method. Theoretical analysis and systematic simulations are conducted to corroborate the effectiveness of the developed estimation method as well as the robustness of the proposed controller.

A Data-Driven Approach to Identify Flight Test Data Suitable to Design Angle of Attack Synthetic Sensor for Flight Control Systems

Digital avionic solutions enable advanced flight control systems to be available also on smaller aircraft. One of the safety-critical segments is the air data system. Innovative architectures allow the use of synthetic sensors that can introduce significant technological and safety advances. The application to aerodynamic angles seems the most promising towards certified applications. In this area, the best procedures concerning the design of synthetic sensors are still an open question within the field. An example is given by the MIDAS project funded in the frame of Clean Sky 2. This paper proposes two data-driven methods that allow to improve performance over the entire flight envelope with particular attention to steady state flight conditions. The training set obtained is considerably undersized with consequent reduction of computational costs. These methods are validated with a real case and they will be used as part of the MIDAS life cycle. The first method, called Data-Driven Identification and Generation of Quasi-Steady States (DIGS), is based on the (i) identification of the lift curve of the aircraft; (ii) augmentation of the training set with artificial flight data points. DIGS’s main aim is to reduce the issue of unbalanced training set. The second method, called Similar Flight Test Data Pruning (SFDP), deals with data reduction based on the isolation of quasi-unique points. Results give an evidence of the validity of the methods for the MIDAS project that can be easily adopted for generic synthetic sensor design for flight control system applications.

Flight envelope estimation for helicopters under icing conditions via the zonotopic reachability analysis

This paper presents a new approach to estimating a flight envelope for a helicopter flying under clean and icing conditions based on a zonotopic reachability analysis. The helicopter is stabilized with a state-feedback linear controller to maintain its attitude around a trim point. Given the sets of possible initial conditions and disturbance inputs, and control effectiveness indices, this analysis results in reachable sets of state variables of the helicopter for a finite time interval. The reachability analysis is thus useful for predicting whether or not the helicopter evolves into an unsafe flight condition when it undergoes degraded control effectiveness ensued from the icing condition. The merit of this approach is shown through a case study about the Raptor-90 unmanned helicopter stabilized with either the linear quadratic regulator (LQR) or the H∞ controller in a hover-flight mode. This helicopter is considered to be affected by both symmetric and asymmetric icing conditions. The case study shows that the H∞ controller outperforms the LQR controller as the former is able to cope with greater loss of control effectiveness caused by the icing conditions. Moreover, as compared to the longitudinal control input, the lateral and directional control inputs are more susceptible to adverse effects of the icing conditions. These results verify not only the effects of ice contamination upon the helicopter dynamics and control, but also the capability of the proposed approach for flight envelope estimation.

Experimental validation of liquid hydrocarbon based fuel rich gas generator for high speed propulsion systems

For long duration hypersonic flight vehicles, research on high speed air-breathing propulsion systems has been progressing across the world for the past few decades. Ramjets/Scramjets fueled with hydrogen/liquid hydrocarbon fuels are the propulsion systems considered for different regimes of flight envelope. Dual Combustion Ramjet (DCR) engine is a prospective candidate for one of such kind of applications, where the engine is expected to work for a wider range of flight Mach number regime. It has a fuel rich subsonic dump combustor and a supersonic combustor, which work in tandem. In the present work, extensive experimental and analytical work has been carried out to develop a fuel rich gas generator often called as dump combustor suitable for operation from 20 to 28 km altitude and in the range of 4–6 flight Mach number. Connect pipe mode testing has been carried out by varying the throat area of dump combustor for Mach 4 conditions. Tests have been carried out with Jet-A fuel at different equivalence ratio and with fuel injection in different combination and stages. Fuel injection scheme used here is of shower head type in intake arm and a single simplex swirler with film cooling holes in the dump combustor. Study has been also carried out by varying the location of swirler along the combustor length. Combustor air entry conditions are simulated using H2–O2 based air heater and a test bed nozzle. An exhaustive amount of tests have been carried out to establish condition suitable for consistent ignition & sustained combustion in this highly fuel rich condition. A One dimensional (1D) mathematical model with NASA CEA interface has been developed to predict & compare the test performance. Sustained ignition is achieved for fuel flow rate of 0.8 kg/s and at an equivalence ratio of 0.8–0.9.

Multi-step predictive compensated intelligent control for aero-engine wireless networked system with random scheduling

Within the future application of wireless network for the aero-engine control problem, resource constraints (caused by the limitation of hardware), network traffic restriction, interference and long time-delay, must be considered as one of the difficulties to be solved, thus, the network connection and transmission efficiency can be ensured. In order to ensure the performance of aero-engine nonlinear wireless networked control system in consideration with complicated flight conditions and adverse network environments, a SVM (Support Vector Machine) intelligent inverse controller with a reference model was established. To reduce the offline modeling error with time-varying random scheduling strategy and disturbance, a LSTM (Long Short Term Memory) online multi-step error compensation was designed. Training result, on the four optimal conditions chosen in the flight envelop, provided the model and controller establishing is completed. After introduced into the overall system, checking result, showed the errors and the output with/without LSTM compensation on the other four random conditions, provided the control performance under the three kind of adverse wireless network environments. for comparison, other intelligent algorithms were also adopted to predict the multi-step errors in the reference model, then the accuracy and computing time provided the advantage of LSTM algorithm. The method and strategy proposed in this paper ensure the aero-engine safe working at the price of some control performance loss on the adverse wireless network environments.

Control Strategy for Helicopter Thermal Management System Based on Liquid Cooling and Vapor Compression Refrigeration

With the continuous application of high-power electronic equipment in aircraft, highly efficient heat transfer technology has been emphasized for airborne applications. In this paper, a thermal management system based on an antifreeze liquid cooling loop and a vapor compression refrigeration loop is presented for high-power airborne equipment in a helicopter. The simulation models of the thermal management system are built in order to study its control strategy for the changing flight conditions. The antifreeze-refrigerant evaporator and air-refrigerant condenser are specially validated with the experimental data. A dual feedforward proportion integration differentiation and expert control algorithm are adopted in the inlet temperature of the cold plate and sub-cooling control of the refrigerant by regulating the compressor speed and the fan speed, respectively. A preheating strategy for antifreeze is set up to decrease its flow resistance in cold day conditions. The control strategy for the thermal management system is finally built based on the above control methods. In this paper, two extreme conditions are discussed, including cold and hot days. Both the simulation results show that the superheated, sub-cooling and antifreeze inlet temperature of the cold plate can be controlled at 3 to8 °C, −10 to −3 °C and 18 to22 °C, respectively. Under the same changing flight envelope, the coefficient of performance of the vapor compression refrigeration loop is relatively stable on the cold day, which is about 6, while it has a range of 2.58–4.9 on the hot day.

Online Safe Flight Envelope Protection for Icing Aircraft Based on Reachability Analysis

Icing encountered by an aircraft in flight can pose a great threat to flight safety, which is the foremost concern in aviation. Ice accretion has nonlinear and coupling properties, and therefore, a conventional envelope protection system cannot successfully deal with the condition. A safe flight envelope based on reachability analysis is proposed in this paper as a basis for designing an online envelope protection system for aircraft under icing conditions. This system uses a neural network method to identify changes in the aerodynamic coefficients and to classify the degrees of icing and uses a database-driven approach to solve the onboard safe flight envelope computation problem. The database of icing safe flight envelopes is created offline and can be retrieved in real time. The Research Civil Aircraft Model (RCAM) encountering icing conditions in the climbing phase was taken as an example to verify the feasibility of the online safe flight envelope protection system. The simulation results showed that the system can classify the degrees of icing efficiently, prevent the aircraft from deviating from the safe flight envelope, guide it back to the envelope along the fastest path, and reduce the risk of loss of control under icing conditions.

A Numerical Investigation of the Geometric Parametrisation of Shock Control Bumps for Transonic Shock Oscillation Control

At transonic flight conditions, shock oscillations on wing surfaces are known to occur and result in degraded aerodynamic performance and handling qualities. This is a purely flow-driven phenomenon, known as transonic buffet, that causes limit cycle oscillations and may present itself within the operational flight envelope. Hence, there is significant research interest in the development of shock control techniques to either stabilise the unsteady flow or raise the boundary onset. This paper explores the efficacy of dynamically activated contour-based shock control bumps within the buffet envelope of the OAT15A aerofoil on transonic flow control numerically through unsteady Reynolds-averaged Navier–Stokes modelling. A parametric evaluation of the geometric variables that define the Hicks–Henne-derived shock control bump will show that bumps of this type lead to a large design space of applicable shapes for buffet suppression. Assessment of the flow field, local to the deployed shock control bump geometries, reveals that control is achieved through a weakening of the rear shock leg, combined with the formation of dual re-circulatory cells within the separated shear-layer. Within this design space, favourable aerodynamic performance can also be achieved. The off-design performance of two optimal shock control bump configurations is explored over the buffet region for M = 0.73, where the designs demonstrate the ability to suppress shock oscillations deep into the buffet envelope.

Evaluating the effects of lateral control surfaces failure on the generic transport model: a case study

ABSTRACTExtensive research in recent years has focused on improving the current loss-of-control prevention systems and developing new strategies for safe path planning of the impaired aircraft. Success in developing such systems requires a comprehensive perception of the influence of damage on the aircraft’s dynamic behaviour and performance, and the effect of various failure degrees on the flight envelope confinement and the remaining safe maneuvers. This paper comprehensively describes the effects of lateral control surface failure on the NASA Generic Transport Model (GTM) flight envelope, defined by a set of attainable steady-state maneuvers herein referred to as trim points. The study utilises a large database of high-fidelity maneuvering flight envelopes computed for the unimpaired case and wide ranges of the aileron and rudder failure cases at different flight conditions. Flight envelope boundary is rigorously investigated, and the key parameters confining the trim points at different boundary sections are identified. Trend analyses of the impaired flight envelopes and the corresponding limiting factors demonstrate the effect of various failure degrees on the remaining feasible trim points. Results can be employed in emergency path planning with potential uses in the development of aircraft resilient control and upset recovery systems.

Modeling, Design, and Flight Testing of Three Flutter Controllers for a Flying-Wing Drone

Three flutter-suppression designs for a flexible flying-wing research drone are discussed, along with the modeling and flight-test results. The drone, with wing span of 10 feet, aspect ratio of almost 9, and control surfaces along the entire wing trailing edges, was developed for flight research into the modeling and control of highly elastic aircraft. The low-speed vehicle was designed to exhibit body-freedom flutter in its flight envelope. The all-important dynamic modeling of the vehicle, used for analysis and control design, is based on a mean-axis formulation and quasi-steady aerodynamics, with the nondimensional aerodynamic/aeroelastic coefficients updated from flight tests. Flutter-modeling results were found to agree with those from NASTRAN and flight tests. The three different flutter-suppression approaches include both fairly classical and multivariable methods, all fixed gains but with different architectures. The primary control-design objective was to augment the damping of the eventual flutter mode at a design condition below the open-loop flutter speed, for safety of flight, and to achieve a closed-loop flutter speed at least as high as the open-loop flutter speed. Analysis and flight tests revealed that all three designs achieved these design goals. In addition, two of these designs actually expanded the flutter boundary. The theoretical stability robustness of all three controllers at the design flight condition is quite good, but there are differences in controller complexity. Research is currently underway to further expand the flutter boundary using all three design approaches.

On the stiffness and damping characteristics of magnetorheological dampers and their implications for flutter analysis

Aeroelastic stability analysis plays an important role during the development of new aircraft. It can affect flight envelope, the aircraft geometry, and the design of components, like control surface actuators, for example. However, the aerial vehicle needs to be designed free of flutter even if the actuators experience failure. For some cases the device can exhibit a dynamic behavior similar to the nonlinear damper. In this context, the present paper discusses the characterization of failed actuators described by a magnetorheological damper (MRD). Linear equivalent parameters of damping and stiffness are extracted from Fourier components at different driven harmonic amplitudes and frequencies of motion. The aeroelastic system stability is evaluated by using an adapted pk-based method. Numerical simulations are carried out in the time domain to demonstrate that this proposed modified pk method allows one to investigate the linear equivalent system to predict the dynamic behavior of the nonlinear system when a MRD is considered.

Design of a Haptic Feedback System for Flight Envelope Protection

Several modern aircraft use a passive control manipulator: a spring–damper system that generates command signals to the flight control computers in combination with a flight envelope protection system that limits pilot inputs when approaching the aircraft limits. This research project aims to increase pilot awareness of this protection system through the use of force feedback on the control device, that is, haptics. This paper describes in detail how the haptic feedback works and when it triggers; another paper will discuss the results of an experimental evaluation. With the current haptic design, pilots can get five cues: first, a discrete force cue when approaching the limits; second, an increased spring coefficient for control deflections that bring the aircraft closer to its limits; third, a stick shaker for low velocities; fourth, if a low-velocity condition requires an input, the stick is moved forward to the desired control input; and finally, the stick follows the automatic Airbus “pitch-up” command during an overspeed condition. This novel system is expected to help pilots correctly assess the situation and decide upon the right control action. It will be evaluated in two scenarios close to the flight envelope limits: a windshear and an icing event.

Wildfire Fighting by Unmanned Aerial System Exploiting Its Time-Varying Mass

This letter presents an approach for accurately dropping a relatively large amount of fire retardant, water or some other extinguishing agent onto a wildfire from an autonomous unmanned aerial vehicle (UAV), in close proximity to the epicenter of the fire. The proposed approach involves a risky maneuver outside of the safe flight envelope of the UAV. This maneuver exploits the expected weight reduction resulting from the release of the payload, enabling the UAV to recover without impacting the terrain. The UAV is tilted to high pitch angles, at which the thrust may be pointed almost horizontally. The vehicle can therefore achieve higher horizontal speeds than would be allowed by conventional motion planners. This high speed allows the UAV to significantly reduce the time spent close to the fire. As a result, the overall high heat exposure is reduced, and the payload can be dropped closer to the target, minimizing its dispersion. A constrained optimal control problem (OCP) is solved taking into account environmental parameters such as wind and terrain gradients, as well as various payload releasing mechanisms. The proposed approach was verified in simulations and in real experiments. Emphasis was put on the real time recalculation of the solution, which will enable future adaptation into a model predictive controller (MPC) scheme.