Steady-state Flight Conditions Research Articles

A minimum objective function trim procedure for VTOLs noise reduction

The present paper proposes a novel approach for multi-rotor acoustic nuisance mitigation based on the identification of optimal control settings. This strategy is possible thanks to the peculiarity of multi-rotor systems (like, for instance, those involved in urban air mobility applications) to have redundant controls. The idea is to take full advantage of the control redundancy by defining the control settings that guarantee the desired steady-state flight conditions while, at the same time, optimizing selected target functions such as performance or noise emissions. To this aim, the trim problem is recast into a constrained cost function minimization problem. Since the objective is to reduce as much as possible the noise disturbance of trimmed VTOL, an aeroacoustic tool based on the chordwise-compact form of the Farassat 1A formulation for the evaluation of noise radiation is suitably coupled with a trim solver. The numerical investigations consider quadcopter and hexacopter configurations in level flight at several advancing speeds. Comparisons of the proposed minimum-noise trim strategy with two other trim strategies based on the minimization of rotor torque and control effort prove its capability to provide an overall reduction of noise throughout the entire velocity range examined.

Optimal Performance Trim Procedure for Multirotor Vertical Takeoff and Landing Vehicles

The present paper proposes a general approach capable of determining the set of trim commands, aerodynamic control surface actuation, and/or engine thrust regulation that allow specific trimmed steady-state flight conditions of nonconventional vertical takeoff and landing vehicles under given performance constraints. Exploiting the peculiarity of these vehicles to have redundant controls, namely more controls than those strictly required to define steady-state flight conditions, their trim problem is recast into a constrained minimization of a cost function. This approach is of particular interest in that it allows the identification of control settings for multirotor vehicles that, at the same time, guarantee trimmed flight conditions and optimization of selected target functions such as, for instance, performance or noise emission. The proposed approach is numerically applied to quadcopter and hexacopter configurations for three flight conditions (level forward flight, climb, and coordinate turn), and three strategies of trim by different trim variables (rotor angular velocity, blade collective pitch, and combination of angular velocity and blade collective pitch). The effect of the introduced cost function is investigated, comparing the outcomes of minimum control effort and minimum torque strategies. The results achieved demonstrate the capability of the proposed algorithm to exploit redundant controls of multirotor vehicles for identifying selected optimal trim operative conditions.

Analysis of the Short-Term Dynamics of Morphing Aircraft Caused by Shape Change Based on the Open-Loop Response and the Reachable Set Theory

This work investigates the short-term dynamics caused by shape changes of morphing aircraft. We select a symmetric variable sweep morphing aircraft as the object of study and establish a six-degree-of-freedom multi-loop cascade model, and the coupling between derivative terms is eliminated by matrix transformation. Considering that the change in aerodynamic shape significantly affects the aerodynamic forces of the aircraft in a short period of time, and the variation in mass distribution generates additional aerodynamic forces and moments, we analyze the effects of these factors on the dynamic characteristics of the aircraft based on the open-loop response starting from the steady-state flight conditions. In addition, we analyze the improvement in maneuvering performance brought by morphing as an additional control input. We apply reachable set theory to multi-loop equations of motion and use the size of the reachable set to measure the maneuverability of aircraft. The results confirm that morphing can effectively improve the maneuverability of the aircraft.

Fan design assessment for BLI propulsion systems

Civil aviation is aiming at fuel efficient aircraft concepts. Propulsion systems using boundary layer ingestion (BLI) are promising to reach this goal. The focus of this study is on the DLR UHBR fan stage of a tube and wing aircraft with rear-integrated engines. In this integration scenario the propulsion system and especially the fan stage receives distorted inflow in steady-state flight conditions. The distortion pattern and distortion intensity are dependent on the operating conditions. Consequently, the interaction of the fan and the distortion changes over the flight envelope. The first part of the paper aims at gaining knowledge of the BLI fan performance in the operating points end of field, approach, cruise (CR) and top of climb (TOC) using high-fidelity, unsteady RANS approaches. The analysis includes fan map performance metrics and a deeper insight into the flow field at CR and TOC. The preliminary design of a fan stage requires fast turn-around times, which are not fulfilled by high-fidelity approaches. Therefore, a fast, throughflow-based methodology is developed, which enables aerodynamicists to design distortion-tolerant fans. The main characteristics of the methodology is outlined in the second part. Consequently, the methodology is taken advantage of to investigate parameter sensitivities in terms of tip speed, blade thickness, solidity, the annulus geometry and a non-axisymmetric stator. This study suggests that distortion-tolerant fans should be designed at higher tip speeds than conventional design experience recommends to limit the local operating point excursion.

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.

F-16/MATV; Optimal Position Controller Design using Robust and Discrete Linear Quadrature Gaussian [RLQG] Controller

This paper presents the robust controller design for nonlinear F-16/MATV (Multi–Axis – Thrust - Vectoring) position control system. The linearization of nonlinear flight position system is guaranteed by feedback linearization using Taylor series expansion and the optimal LQG controllers are designed to achieve the steady state flight condition. The comparison of the using continuous and discrete LQG is fully discussed for control system. First the continuous optimal LQG controller is designed for continuous time state space represented flight dynamic system based on separation principle. But the designed optimal LQG controller requires a very large controller gain to achieve the design objective, which lacks high sensitivity and leads to highly cost system. This leads to design a controller using discrete LQG controller for discrete time state space represented flight control system. After implementing this controller into the system in MATLAB Simulink environment, using singular value decomposition technique, it is founded that the performance and robustness of the design objective is satisfied.

Thermal–structural analysis of the platelet heat-pipe-cooled leading edge of hypersonic vehicle

One of the main challenges for the hypersonic vehicle is its thermal protection, more specifically, the cooling of its leading edge. To investigate the feasibility of a platelet heat-pipe-cooled leading edge structure, thermal/stress distributions for steady-state flight conditions are calculated numerically. Studies are carried on for IN718/Na, C-103/Na and T-111/Li compatible material combinations of heat pipe under nominal operations and a central heat pipe failure cases, and the influence of wall thickness on the design robustness is also investigated. And the heat transfer limits (the sonic limit, the capillary limit and the boiling limit) are also computed to check the operation of platelet heat pipes. The results indicate that, with a 15mm leading edge radius and a wall thickness of 0.5mm, C-103/Na and T-111/Li combinations of heat pipe is capable of withstanding both nominal and failure conditions for Mach 8 and Mach 10 flight respectively.

Modelling of a Small Unmanned Aerial Vehicle

Unmanned aircraft systems (UAS) are playing increasingly prominent roles in defense programs and defense strategies around the world. Technology advancements have enabled the development of it to do many excellent jobs as reconnaissance, surveillance, battle fighters, and communications relays. Simulating a small unmanned aerial vehicle (SUAV) dynamics and analyzing its behavior at the preflight stage is too important and more efficient. The first step in the UAV design is the mathematical modeling of the nonlinear equations of motion. A survey with a standard method to obtain the full non-linear equations of motion is utilized, and then the linearization of the equations according to a steady state flight condition (trimming) is derived. This modeling technique is applied to an Ultrastick-25e fixed wing UAV to obtain the valued linear longitudinal and lateral models. At the end, the behavior of the states of the non-linear UAV and the resulted linear model is compared by applying a doublet signals in the control surfaces to check the matching between them.

Numerical Investigation on Actively Cooled Thermal Protection Systems with Ni-Based Alloys

One of the greatest challenges of hypersonic vehicles is their thermal protection and, more specifically, the cooling of their engine. To simulate the behavior of a complete actively cooled thermal protection system, a computational fluid dynamics and finite element analysis coupling method is applied to calculate the fluid/thermal/stress distributions for steady-state flight conditions. Work has been done on four different Ni-based alloys and three different panel structures. Temperature and stress profiles at the outlet cross section show that the maximum temperature and stress happen on the side that is close to the combustion chamber, and so this is the section on which the active cooling system should focus. It is better to have small rounded chamfers in the panels to decrease the stress concentration at the corners. Failure maps are presented for four Ni-based alloys showing the comparison of their thermostructural performance, which will be helpful for the selection of the materials in an active cooling system.

Nonlinear $\boldsymbol{\mathcal{H}}_{\boldsymbol\infty}$ Control of UAVs for Collision Avoidance in Gusty Environments

This paper proposes a nonlinear $\mathcal{H}_{\infty}$ controller for stabilization of velocities, attitudes and angular rates of a fixed-wing unmanned aerial vehicle (UAV) in a windy environment. The suggested controller aims to achieve a steady-state flight condition in the presence of wind gusts such that the host UAV can be maneuvered to avoid collision with other UAVs during cruise flight with safety guarantees. This paper begins with building a proper model capturing flight aerodynamics of UAVs. Then a nonlinear controller is developed with gust attenuation and rapid response properties. Simulations are conducted for the Shadow UAV to verify performance of the proposed controller. Comparative studies with the proportional-integral-derivative (PID) controllers demonstrate that the proposed controller exhibits great performance improvement in a gusty environment, making it suitable for integration into the design of flight control systems for cruise flight of UAVs.

Fluid-Structure Coupled Analysis of a Transport Aircraft and Flight-Test Validation

The estimation of aerodynamic loads of a transport aircraft taking into account static aero-elastic deformations at steady-state flight conditions is described. Within the scope of a pilot study a fluid-structure coupled (FSC) simulation approach based on state of the art numerical fluid dynamics and structural analysis methods has been applied to an Airbus A340-300 aircraft in both cruise and high-lift configurations. Coupled analyses were performed using an in-house simulation procedure linking DLR’s flow solver TAU and the commercial finite-element code NASTRAN™. Numerical results were validated against flight test data generated in the research project AWIATOR (Aircraft Wing with Advanced Technology Operation) funded by the European Union. A good agreement of simulation results and flight test data, including both fluid dynamics and structural deformation properties, was observed. Aero-elastic effects in cruise flight were found to be larger than in high-lift, where only slats and the wing’s leading edge region are affected. Work presented in this paper is part of a co-operation between DLR and Airbus Deutschland within the joint research project HIT (High-Lift Innovative Technologies).

Fuzzy logic‐based detection scheme for pilot fatigue

PurposeThe paper aims to present the development of a detection scheme for pilot fatigue using fuzzy logic. Evaluation parameters based on the dynamic response of the pilot/aircraft system are to be defined and criteria for online fatigue detection to be formulated.Design/methodology/approachThe approach is based on the idea that, while performing the same task, under otherwise identical conditions, the dynamic signatures of the pilot/aircraft system are different depending on the pilot condition, “rested” or “tired.” Tests performed on a 6 degrees‐of‐freedom (DOF) flight simulator with pilots at two extreme levels of alertness are used to define parameters based on aircraft states and pilot input measurements that can serve as pilot fatigue detectors at steady state flight conditions. These parameters are computed using the statistics of the tracking errors (TE), state and control time histories, and the Fourier transforms of the TE. Fuzzy logic is used to evaluate the pilot condition based on composite detection parameters.FindingsValidation tests on a 6 DOF flight simulator showed that the proposed detection scheme has promising capabilities for safety monitoring purposes and design of control laws that can accommodate for pilot abnormal conditions.Research limitations/implicationsThe pilot fatigue detection algorithm presented in this paper can be used as a starting point for future research in the following directions: development of safety monitoring systems for warning and/or triggering of automatic control compensation; development of pilot fault‐tolerant control laws; development of human pilot models for simulation, handling qualities assessment, and control laws design.Practical implicationsThe approach for pilot fatigue detection proposed in this paper is a viable alternative to existing methods based on physiological measurements such as electrical activity of the brain, pulse, body temperature, etc. which imply direct and permanent connection of the pilot to the measurement system and interfere adversely with pilot comfort and his/her ability to perform the task. The proposed approach eliminates this drawback and does not require on‐board additional heavy equipment.Originality/valuePilot fatigue assessment from measurements of pilot/aircraft dynamic parameters has not yet been investigated as an alternative to the physiological approach.

A Materials Selection Protocol for Lightweight Actively Cooled Panels

This article provides a materials selection methodology applicable to lightweight actively cooled panels, particularly suitable for the most demanding aerospace applications. The key ingredient is the development of a code that can be used to establish the capabilities and deficiencies of existing panel designs and direct the development of advanced materials. The code is illustrated for a fuel-cooled combustor liner of a hypersonic vehicle, optimized for minimum weight subject to four primary design constraints (on stress, temperatures, and pressure drop). Failure maps are presented for a number of candidate high-temperature metallic alloys and ceramic composites, allowing direct comparison of their thermostructural performance. Results for a Mach 7 vehicle under steady-state flight conditions and stoichiometric fuel combustion reveal that, while C–SiC satisfies the design requirements at minimum weight, the Nb alloy Cb752 and the Ni alloy Inconel X-750 are also viable candidates, albeit at about twice the weight. Under the most severe heat loads (arising from heat spikes in the combustor), only Cb752 remains viable. This result, combined with robustness benefits and fabrication facility, emphasizes the potential of this alloy for scramjets.

Testing coupled rotor blade–lag damper vibration using real-time dynamic substructuring

In this paper, we present new results from laboratory tests of a helicopter rotor blade coupled with a lag damper from the EH101 helicopter. Previous modelling of this combined system has been purely numerical. However, this has proved challenging due to the nonlinear behaviour of the dampers involved—the fluid filled lag damper is known to have approximate piecewise linear force–velocity characteristics, due to blow-off valves which are triggered at a certain force level, combined with a strongly hysteretic dynamic profile. The novelty of the results presented here, is that the use of a hybrid numerical–experimental testing technique called real-time dynamic substructuring, allowed a numerical model of the rotor to be combined with the physical testing of a flight certified lag damper unit. These hybrid tests, which are similar in concept to hardware-in-the-loop, were carried out in real-time such that there is bi-directional coupling between the numerical blade model and the experimental lag damper. The new results obtained from these tests (for steady-state flight conditions) reveal how the inclusion of a real damper produces a more realistic representation of the dynamic characteristics of the overall blade system (during operational flight conditions) than numerical modelling alone.

Investigation on the flight characteristics of a conceptual fluidic thrust-vectored aerial tail-sitter

The feasibility of integrating co-flow fluidic thrust-vectoring idea into the dynamics of a small flapless aerial tail-sitter is investigated in this article. The aircraft trimmability in different phases of flight and stability in take-off and level flight, are the main issues of concern for the study presented herein. In this respect, the vehicle's characteristic equations are derived by linearization of the general non-linear equations of motion. Since the vehicle was supposed to be merely controlled by fluidic thrust-vectoring, the concept was novel and some new derivatives are introduced. Margins of the required thrust-vector angle, to obtain a steady-state flight condition, are provided by use of the trim analysis. In addition, stability in take-off transition and level flight is considered utilizing the values of the system poles. Non-linear simulation is developed to examine the aircraft responses to the command inputs and to generate the aircraft trajectory. It is found that the integrated fluidic thrust-vectoring system can be employed as a viable means to achieve both conventional and transition flight for the proposed vehicle. The obtained mathematical model reveals the dependence of the aircraft performance and design criteria on the presented method. This model will be used for further studies on the general aspects of flight dynamics and control systems design for such vehicle.

Design and performance evaluation of residual genertors for the FDI of an aircraft

In this work, several procedures for the fault detection and isolation (FDI) on general aviation aircraft sensors are presented. In order to provide a comprehensive wide-spectrum treatment, both linear and nonlinear, model-based and data-driven methodologies are considered. The main contributions of the paper are related to the development of both FDI polynomial method (PM) and FDI scheme based on the nonLinear geometric approach (NLGA). As to the PM, the obtained results highlight a good trade-off between solution complexity and resulting performances. Moreover, the proposed PM is especially useful when robust solutions are required for minimising the effects of modelling errors and noise, while maximising fault sensitivity. As to the NLGA, the proposed work is the first development and robust application of the NLGA to an aircraft model in flight conditions characterised by tight-coupled longitudinal and lateral dynamics. In order to verify the robustness of the residual generators related to the previous FDI techniques, the simulation results adopt a typical aircraft reference trajectory embedding several steady-state flight conditions, such as straight flight phases and coordinated turns. Moreover, the simulations are performed in the presence of both measurement and modelling errors. Finally, extensive simulations are used for assessing the overall capabilities of the developed FDI schemes and a comparison with neural networks (NN) and unknown input Kalman filter (UIKF) diagnosis methods is performed.

Effect of flapping trajectories on the dragonfly aerodynamics

The effects of translational, figure-eight and double-figure-eight flapping trajectories on the dragonfly aerodynamics were numerically studied by solving the Navier-Stokes equations. There is a common characteristic regarding the lift/drag force coefficients that the downstroke flapping provides the lift forces while the upstroke flapping creates the thrust forces for different flapping trajectories. The maximum lift force coefficient exceeds five for the translational trajectory. It is greater than six for the figure-eight and double-figure-eight flapping trajecto- ries, which is sufficiently larger than unity under the steady state flight condition. The ellipse and dou- ble-figure-eight flapping trajectories yield the de- crease of the lift force, while the figure-eight flapping trajectory yields higher lift force as well as the thrust force than the translational flapping one. During the insect flight, the wing flapping status should be changed instantaneously to satisfy various require- ments. Study of the flapping trajectories on the insect aerodynamics is helpful for the design of the Mi- cro-air-vehicles (MAVs).

Wind-tunnel evaluation of mid-airframe installed turbojet engines

A series of power-on wind-tunnel tests have been conducted to evaluate the installed performance of midairframe mounted tactical-missile turbojet sustainers. Two low-cost, expendable turbojet engine designs were installed in a high-fidelity wind-tunnel model of a specific tactical missile. To accommodate the unique installation requirements, each engine design incorporated bifurcated pitot inlets and side-exhausting bifurcated exhaust ducts. Each installed engine was fully functional and was intended to closely replicate the in-flight operational response. Power-on wind-tunnel evaluations were conducted for each engine model that encompassed a wide range of sea level, steady-state flight conditions. The test condition variables evaluated were Mach number, pitch angle, yaw angle, engine throttle setting, and control fin deflection. Missile axial force data were acquired to determine installed, delivered engine net thrust. In addition, extensive thermal instrumentation was installed on the model to evaluate the influence of the turbojet exhaust plume on the aft section of the missile. A detailed description of the test program is provided. Detailed descriptions of the hard ware configuration and test variables are presented. Presented are discussions of theoretical engine performance models, experimental installed performance results, projections of overall missile performance, and evaluations of the effect of pitch on engine performance. A detailed evaluation of the thermal impact of turbojet operation is provided. Discussions of the thermal impact of pitch, yaw, and fin deflection angles are presented. The results presented demonstrate that both sustainer configurations, successfully operated in an installed configuration under in-flight conditions, delivered adequate installed performance to satisfy missile system requirements, and had minimal adverse thermal impact on the airframe. The results of the evaluation fully verify the viability of employing turbojet engines in a mid-airframe installation.

Aircraft Control System Design and Analysis Program Package

A description will be given of a software package for the design and analysis of aircraft control systems. The package offers the possibility of modelling, optimization and evaluation of controllers for deterministic or stochastic, time-continuous or discrete-time systems. According to the phases of the design process, the software package is divided into the following four sections: 1. calculation of a steady-state flight condition and specification of the linearized equations of motion; 2. compilation of subsystems to obtain the total system model; 3. application of linear system theory to computation of observer and controller gains; 4. computation and evaluation of the responses of the controlled aircraft. Response evaluations may be extended using a nonlinear flight simulation program, which may include either simulation of a digital control system or simulation of the human pilot controlling the aircraft with the aid of a so-called flight-director system.

Turnover of locust haemolymph diglycerides during flight and rest

The constant, elevated diglyceride concentration in the haemolymph of adult male Locusta migratoria during sustained tethered flight (14.1 ± 1.2 mg/ml haemolymph) was established to be a steady state, in which the rate of diglyceride mobilization matched the rate of utilization. Pulse-labelling of the haemolymph diglycerides with [1- 14C]-oleic acid in steady state flight conditions results in a turnover time of about 1 hr for the diglyceride pool, whereas in animals rested after flight values close to 8 hr were obtained. So, during flight there is a sharp increase in lipid transport, the initial turnover rate in the resting animal (0.3–0.5 mg diglyceride/hr) reaching 3.4 mg diglyceride/hr at the steady state flight level. As identical turnover rates were obtained using [1- 14C]-palmitic acid or [1- 14C]-linolenic acid, apparently there is no preferential utilization of diglycerides with a specific fatty acid composition in flight muscle energy metabolism. The characteristic diglyceride fatty acid spectrum at the steady state flight level remained remarkably constant during flight performances lasting for 6 hr. The above findings indicate homogeneity of the haemolymph diglyceride pool and suggest random utilization of fatty acids.