Wide Range Of Flight Conditions Research Articles

Aeroelastic predictions for steady and unsteady flow characteristics of the HIRENASD wing

In this study, we focus on static and dynamic aeroelastic analyses of the HIRENASD wing based on reference experimental data for two different flight conditions for Aeroelastic Prediction Workshop-1. The major anticipations from the HIRENASD project are to improve the knowledge about aero-structural dynamics, and to get experimental and computational data in a wide range of flight conditions. The experiments have been formerly conducted in cryogenic medium to investigate steady and unsteady aeroelastic responses in transonic regime for low and high Reynolds numbers. For current aeroelastic computations, first, a free vibration analysis is performed in Nastran using the latest structural model provided by NASA. Sequentially, the modal solution is imported from the finite element solver to Zeus software which is an Euler equations based aeroelastic solver coupled with integral boundary layer method. The aerodynamic model and fluid structure interaction parameters are constructed in Zeus. Steady and unsteady aeroelastic results are examined for the specified stations along the wing span, and interpolated for chordwise direction so as to match them with the wind tunnel test points. The designated output parameters such as steady aerodynamic lift, moment and drag coefficients, and steady and unsteady pressure distributions along the chordwise direction are compared to the experimental data and NASA's FUN3D code results which are provided in literature.

DLR natural and hybrid transonic laminar wing design incorporating new methodologies

AbstractIn the present work natural laminar flow (NLF) and hybrid laminar flow (HLF) wing designs are presented which were obtained by combining new methodologies with experience and knowledge obtained with traditional laminar wing design methods. The NLF wing design is performed for wing-body configurations with backward swept wing (BSW) and forward swept wing (FSW). Initial aerofoil sections were obtained by using a new sectional conical wing method which allows the design of transonic wing sections, taking into account the effects of sweep and taper for the computational cost of a 2D method. Except for flow regions with strong 3D influence, wings constructed with these aerofoils showed an acceptably large region with laminar boundary layer and small shocks at design and specified off-design conditions. For regions close to the body and the tip a 3D inverse design method was further required. For the BSW case, due to cross flow a premature transition occurred. Therefore, a HLF panel was required to obtain a larger laminar region. A suction distribution was obtained using the suction distribution module of the automated target pressure generator (ATPG). This generator optimises the pressure distributions in terms of minimising drag while keeping certain boundary conditions constant, e.g. lift and momentum. Using the ATPG, the laminar extent of the BSW NLF design could be further improved for the inboard wing. With the new methodologies design work was reduced. They lead to a design with reserves that allow for acceptable off-design performance qualities by keeping the wing laminar over a wide range of flight conditions.

Engine performance analysis and optimization of a dual-mode scramjet with varied inlet conditions

A dual-mode scramjet can operate in a wide range of flight conditions. Higher thrust can be generated by adopting suitable combustion modes. Based on the net thrust, an analysis and preliminary optimal design of a kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6 were investigated by using a modified quasi-one-dimensional method and simulated annealing strategy. Engine structure and heat release distributions, affecting the engine thrust, were chosen as analytical parameters for varied inlet conditions (isolator entrance Mach number: 1.5–3.5). Results show that different optimal heat release distributions and structural conditions can be obtained at five different inlet conditions. The highest net thrust of the parameterized dual-mode engine can be achieved by a subsonic combustion mode at an isolator entrance Mach number of 2.5. Additionally, the effects of heat release and scramjet structure on net thrust have been discussed. The present results and the developed analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets. A dual-mode scramjet can achieve high thrust in a wide range of flight conditions. The present work uses modified quasi-1-D analysis model and simulated annealing strategy to investigate the engine performance of kerosene-fueled parameterized dual-mode scramjet at a crucial flight Mach number of 6. Engine structure and heat release distributions were optimized to obtain optimal engine net thrust for varied inlet conditions. The highest net thrust can be achieved by subsonic combustion mode at an isolator entrance Mach number of 2.5. The present results and analytical method can provide guidance for the design and optimization of high-performance dual-mode scramjets.

Aerodynamic optimisation of a camber morphing aerofoil

An aircraft that has been carefully optimised for a single flight condition will tend to perform poorly at other flight conditions. For aircraft such as long-haul airliners, this is not necessarily a problem, since the cruise condition so heavily dominates a typical mission. However, other aircraft such as UAVs, may be expected to perform well at a wide range of flight conditions. Morphing systems may be a solution to this problem, as they allow the aircraft to adapt its shape to produce optimum performance at each flight condition. Optimisation of morphing aerofoils is typically performed separately to the morphing mechanism design. In this work, an optimisation strategy is developed to account for a known possible morphing system within the aerodynamic optimisation process itself. This allows for the limitations of the system to be considered from the start of the design process. The Fishbone Active Camber (FishBAC) camber morphing system is chosen as the example mechanism, and it is shown that the FishBAC can achieve large improvements in performance over non-morphing aerofoils when multiple flight conditions are considered. Additionally, its performance is compared to an aerofoil whose shape can change arbitrarily (as if a perfect morphing mechanism can be designed), and it is shown that the FishBAC performs nearly as well, despite being a relatively simple mechanism.

A Numerical Study on Smart Material Selection for Flapped and Twisted Morphing Wing Configurations

The developments of innovative adaptive structures on Unmanned Aerial Vehicles (UAVs), such as morphing wings, can potentially reduce system complexities by eliminating control surfaces and their auxiliary equipment. This technology has the potential of allowing a UAV to adapt to different mission requirements or to execute a particular mission more effectively by maintaining an optimum airfoil section over a range of speeds for different segments of a mission profile. Studies on a number of smart materials candidates are currently available in the open literature to achieve wing morphing. The material selection depends on several factors including fast dynamic response, low weight, capability to operate over a wide range of flight conditions and low power consumption. This paper presents a review on smart materials technologies for UAV morphing wings. A numerical study in terms of power requirements is also presented for two morphing wing concepts: flapped and twisted wing planforms. The energy calculations for both morphing configurations were based on a two-step procedure. The first step consists of computing the aerodynamic energy using an in-house Vortex-Lattice (VL) based program. Subsequently the pressure field obtained from the first step is then mapped into a finite element mesh and the structural strain energy is calculated. The numerical results indicated that flapped morphing wings have a better aerodynamic performance when compared to twisted wings and different morphing levels can be achieved using lighter smart materials with lower specific energy for this configuration.

Linear-Frequency-Domain Predictions of Dynamic-Response Data for Viscous Transonic Flows

Determining the flutter boundaries for full aircraft configurations by time-accurately solving the Reynolds-averaged Navier–Stokes equations is prohibitive with respect to computational expense, as the unsteady aerodynamic loading must be predicted for a wide range of flight conditions, frequencies, and structural mode shapes. Nonetheless, there is an increasing demand to accurately predict flutter boundaries in the viscous transonic regime—a demand, which, until recently, could only be satisfied by high-fidelity Reynolds-averaged Navier–Stokes methods. Brought to application readiness over the last years, time-linearized/small-disturbance methods, however, have been shown to satisfy this demand as well. They retain the Reynolds-averaged Navier–Stokes method’s fidelity to a high degree, at a substantially reduced computational expense. Such a method is presented here on the basis of the TAU–Reynolds-averaged Navier–Stokes method. Denoted as the TAU linear-frequency-domain method, it is validated for both a standard transonic airfoil and a high-aspect-ratio-wing dynamic test case using rigid pitch modes. The response data obtained from the linear frequency domain are in good agreement with the experiment for a two-dimensional case. For the three-dimensional case, there are larger differences. More important, the linear-frequency-domain method is in excellent agreement to time-accurate Reynolds-averaged Navier–Stokes simulations. Depending on the linear-frequency-domain-employed-solution scheme, reductions in computational costs well beyond an order of magnitude are obtained. In addition, the limits of the so-called frozen-eddy-viscosity approach are established.

Statistical analysis of Airborne Aero-Optical Laboratory optical wavefront measurements

The Airborne Aero-Optical Laboratory has produced a large database of aero-optical measurements with a high-speed, high-resolution Shack Hartmann wavefront sensor. The data have been collected over a wide range of flight conditions. An analysis of the statistical characteristics of the subsonic and early transonic data is performed to assess the adequacy of the spatial and temporal resolution of the data. Sample rate requirements for a minimum variance phase estimator are also explored. The techniques employed are validated by application to mea- surements of optical atmospheric turbulence where results can be antici- pated based on established Kolmogorov statistics. © The Authors. Published by SPIE under a Creative Commons Attribution 3.0 Unported License. Distribution or reproduction of this work in whole or in part requires full attribution of the original publication, including its DOI. (DOI: 10.1117/1.OE.52.7.071416)

Validation of Sliding Mode Observer FDI Schemes on the ADDSAFE Functional Engineering Simulator

This paper presents evaluation results from using second order sliding mode observer FDI schemes on the ADDSAFE Functional Engineering Simulator. This has provided fault detection and isolation design verification and performance assessment in preparation for the industrial evaluation in DEIMOS. This paper presents design verifications for three ADDSAFE benchmark problems: one sensor fault scenario and two actuator fault scenarios. The performance of each of the FDD schemes is assessed using the FES's pseudo-Monte-Carlo setup, where a predetermined grid or sequence is used to cover a wide range of flight conditions and perturbations in the aircraft parameters. Detection logic and statistical assessment for the three FDD schemes are presented.

Digital adaptive control of a winged rocket applicable to abort flight

Since reusable launch vehicles (RLVs) operate in a wide range of flight conditions, the values of the parameters of the dynamic equations are not constant. Therefore, some adaptive control methods for RLVs have been proposed, and digital control systems which are suitable for digital computers. However, the control performance decreases when the nonlinearity strengthens, although linear adaptive control gives an excellent performance when the nonlinearity of the control system can be disregarded. In this article, we apply a digital adaptive feedback linearization control method with time-scale separation to a winged rocket in an abort case. The simulation results show the effectiveness of the control system.

Launch vehicle attitude control using sliding mode control and observation techniques

In determining flight controls for launch vehicle systems, several uncertain factors must be taken into account, including a variety of payloads, a wide range of flight conditions and different mission profiles, wind disturbances and plant uncertainties. Crewed vehicles must adhere to human rating requirements, which limit the angular rates. Sliding mode control algorithms that are inherently robust to external disturbances and plant uncertainties are very good candidates for improving the robustness and accuracy of the flight control systems. Recently emerging Higher Order Sliding Mode (HOSM) control is even more powerful than the classical Sliding Mode Controls (SMC), including the capability to handle systems with arbitrary relative degree. This paper proposes sliding mode launch vehicle flight controls using classical SMC driven by the sliding mode disturbance observer (SMDO) and higher-order multiple and single loop designs. A case study on the SLV-X Launch Vehicle studied under a joint DARPA/Air Force program called the Force Application and Launch from CONtinental United States (FALCON) program is shown. The intensive simulations demonstrate efficacy of the proposed HOSM and SMC-SMDO control algorithms for launch vehicle attitude control.

Aerodynamics of aerofoil sections measured on a free-flying bird

Birds are adapted to a wide range of flight conditions, from steady fixed-wing glides to high angle of attack manoeuvres involving unsteady separated flows. They naturally control and exploit the transitional Reynolds number regime of Re≈ 105 that is currently of interest in unmanned air vehicle technologies. This article presents a reconstruction of the inner portion of a wing of an eagle in free flight, during a rapid pitch-up manoeuvre at the end of a shallow glide to an elevated perch. Photogrammetric techniques were used to map the identified points on the wing and these were used to fit a mathematical model of the upper and lower surface topography using polynomial regression techniques. The surface model accounts for spanwise twist, spanwise bending, and varying chord distribution, as well as for the shape of the aerofoil. The aerodynamics of the two-dimensional aerofoil sections were analysed using XFOIL and were compared against two technical aerofoils, namely the Selig S1223 and Clark Y aerofoils, at 1×105≤ Re≤2×105. The bird aerofoil maintains a robust, near-constant drag coefficient over a wide lift coefficient range.

Investigation of High-Speed Ramjet Engines

This paper gives a brief overview of investigations and developments of combustion chambers of high-speed ramjet engines performed at Moscow Aviation Institute (MAI). Shadow photographs are given which show changes in the structure of regular shocks versus heat input to the flow during hydrogen combustion. The impact of oblique shock waves on combustion is studied by visualizing OH radical luminescence. Because of the significant dependence of the combustion effectiveness on the flow parameters and structure, it is relevant to study ramjet engine designs with a variable air flow duct of the combustion chamber, which allows adjustment of the working section to obtain the required parameters and characteristics over a wide range of flight conditions. A large series of tests of such combustion chambers have been performed under a joint French—Russian program. Measurement results corresponding to flight Mach numbers 2–7 are given. Mixtures of chemically pure methane and hydrogen was used as fuel. Characteristics of a three-stage heater of model flow developed at MAI for testing wide-range combustion chambers are presented.

Identification of nonlinear aerodynamic derivatives using classical and extended local model networks

Determining aerodynamic models for use in simulators requires the model to be valid over a wide range of flight conditions. Local model networks are suitable for this kind of task because they build a global model through a weighted superposition of local simple models. The location of the local models, i.e. the partitioning into submodels is determined automatically as part of the algorithm. Unlike neural networks that yield only black-box models, the structure and parameters of local model networks are interpretable and can quite easily be transformed into modeling functions or table models. Using flight test data, it is shown that local model networks are useful in the identification of models that have to cover a broad range of flight conditions. When identifying aerodynamic parameters from flight test data, often the task is to derive models for the different nonlinear derivatives directly from measurements of the overall coefficient. For this, two extensions of the classical local model networks are introduced and investigated. Out of the two approaches, the structured local networks yield very promising results.

Aerobasics — An introduction to aeronautics

Stability and controllability of an airplane are essential qualities required for safely and effectively performing any flight. In this article we present the concepts of static and dynamic stability and their relation to controllability of an airplane. This leads us to the concept of handling qualities which are expressed in terms of the properties of transient motions of an airplane following a disturbance. It is difficult to design an airplane for satisfactory handling qualities over a wide range of flight conditions (altitude, flight speed, weight and centre of gravity). It is often advantageous to enhance the flying qualities by using an active control technology. Currently, this technology is implemented on many airplanes by using a fly-by-wire control system. A brief description of this type of control system and its advantages are presented.

Some Methods of High Handling Qualities Providing in Whole Flight Envelope

AbstractThe energy approach to flight control has shown the high efficiency. The rational structure of energy control system at landing approach has been created on its basis. The structure of a modified energy flight control system for aircraft with essential non-linear aerodynamics has been developed. However, the system with fixed coefficients couldn't provide the acceptable control quality for maneuverable high-speed aircrafts over the whole flight envelope. Therefore the adjustment of energy control system parameters was needed. For this purpose the new version of a modal method of system design with the required quality has been developed. The optimum system parameters for the operational area were found with its help. The loop of coefficients adjustment was entered into system structure. Modeling has shown invariance of handling qualities in a wide range of flight conditions

Numerical simulation of magnus force control for projectiles configurations

Present paper proposes, through numerical simulations, some optimization possibilities for an innovative concept aimed to control the magnus effect induced by the high spinning velocity of a Secant–Ogive–Cylinder–BoatTail projectile. It appears that with a relatively small contra-spinning surface, the lateral force can be drastically reduced on a wide range of flight conditions. Present computations confirm experiments in supersonics and extend them for high subsonic and transonic regimes for small angles of attack.

Overview of the DARPA Smart Wing Project

The recently completed DARPA/AFRL/NASA Smart Wing program, performed by a team led by Northrop Grumman Corporation (NGC), addressed the development and demonstration of smart materials based concepts to improve the aerodynamic and aeroelastic performance of military aircraft. This paper presents an overview of the program. The program was divided into two phases. Under Phase 1 (January 1995 to February 1999), the NGC-led team developed adaptive wing structures with integrated actuation mechanisms to replace standard hinged control surfaces and provide variable, optimal aerodynamic shapes for a variety of flight regimes. An important limitation of the Phase 1 effort, the low bandwidth achievable in Shape Memory Alloy (SMA)-based actuation, was addressed in Phase 2 (January 1997 to November 2001). Under Phase 2, a 30-percent scale full span wind tunnel model of an NGC Uninhabited Combat Air Vehicle (UCAV) design was developed. For the Phase 2 first wind tunnel test, completed in March 2000, SMA-actuated, hingeless, smoothly contoured, flexible leading and trailing edge control surfaces were incorporated on one wing of the model and conventional trailing edge control surfaces actuated using electric motors on the other. This test provided baseline steady-state data at Mach numbers ranging from 0.3 to 0.8. The test also demonstrated the benefits of the smart leading edge control surface to compensate for loss of aileron effectiveness with increasing dynamic pressure. The second test, performed in May 2001, demonstrated a hingeless, smoothly contoured, structurally compliant, trailing edge control surface actuated using piezoelectric motors. Spanwise and chordwise shape control was demonstrated for the smart trailing edge control surface at deflection rates of up to 80°/s. Performance improvements in terms of increased rolling and pitching moments and lower control surface deflections were quantified. The work performed under this program has demonstrated the feasibility of developing smart control surface designs to provide optimal aerodynamic performance at a wide range of flight conditions.

Bypass-Regulated Solid Fuel Ramjet Combustor in Variable Flight Conditions

A comprehensive experimental investigation aimed at the examination of a solid fuel ramjet (SFRJ) regulation concept using an air-division valve has been conducted. By controlling the ratio between the port and bypass flows, one can maintain a desirable working state over a wide range of flight conditions. A general regulation law based on maintaining a constant fuel-to-air ratio has been tested. The experimental strategy was to isolate key parameters influencing the fuel regression rate, testing them one at a time, keeping the other parameters constant. Experiments aimed at airflow rate, total air temperature and pressure, as well as port diameter effects, simulated a broad flight envelope, from Mach 1.5 to Mach 4.7 and from sea level to 13-km altitude. The test results showed a very good agreement with the theory, demonstrating the feasibility, effectiveness, and characteristics of the air-division valve regulation technique in SFRJ motors.

Future challenges for powerplant aerodynamic integration in combat aircraft

AbstractThe operational requirements of modern combat aircraft demand complex engine intake and exhaust systems, capable of working efficiently over a very wide range of flight conditions and throttle settings. In addition to high aerodynamic efficiency and avoidance of high distortion levels at the engine face, these systems must also meet rigorous radar and infra-red signature targets. This paper discusses the implications from the aerodynamics point of view. Examples of technical approaches which seek to balance the sometimes conflicting requirements of aerodynamics and signatures are outlined. The potential offered by in-flight thrust vectoring to enhance flight performance and/or safety is also reviewed and the aerodynamic implications considered. Overall, propulsion integration for combat aircraft presents several challenges to the aerodynamicist, not least the development and validation of improved theoretical design methods capable of analysing the highly complex flows involved.

New particle formation in aircraft exhaust plumes

We have developed a unique representation of the properties of volatile particles produced by jet aircraft. Our new parameterisation is based upon detailed microphysical mechanisms that incorporate the latest phenomenology observed in aircraft plumes, including the roles of chemi-ions and organic vapours. By comparing particle measurements taken under a wide range of flight conditions, and normalising these using our model, we have reduced dramatically the large residual variability in the volatile particle emission indices inferred from field observations. By coupling our model to a global tracer model, we have been able to narrow considerably the range of predicted global enhancements in background aerosol surface area densities associated with commercial aircraft operations (not discussed here). Our novel approach should prove to be useful in future evaluations of the atmospheric perturbations associated with aviation.