Wide Range Of Flight Conditions Research Articles

Controlling hypersonic boundary layer transition with localized cooling and metasurface treatments

This study investigates a novel method to control hypersonic boundary layer transition using a combined local cooling and local metasurface treatment. The method’s effectiveness was investigated on a 5-degree half-angle blunt wedge with a nose radius of 0.0254 mm at a freestream Mach number of 6.0 using direct numerical simulations and linear stability theory. We explored four cases: (i) adiabatic baseline case, (ii) locally cooled case, (iii) local metasurface case, and (iv) combined local cooling-local metasurface case. Results showed that the combined local cooling-local metasurface treatment significantly reduced both wall pressure disturbance amplitude and the density perturbation amplitude around the sonic line, indicating a potential for controlling hypersonic boundary layer transition. In the local cooling-local metasurface case, the disturbance amplitude at the end of the computational domain was 270 times lower than the baseline case. The study also examined the impact of Reynolds numbers, ranging from 25.59 million per meter to 32.80 million per meter. Unsteady simulations revealed that the Reynolds number had a negligible effect on the local cooling-local metasurface performance, indicating that the proposed method applies to a wide range of flight conditions.

Effects Of Wing Elasticity On Tip Vortex: 3D Deformation And 2D3C Flow Field Measurements

As aircraft embrace broader applications, the flexible wing presents a promising avenue to enhance aircraft performance across a wide range of flight conditions. Although the aerodynamic benefits of wing deformation are widely recognized, the present understanding of the tip vortex is relatively limited. This work experimentally examines the effects of wing deformation on the tip vortex of a simplified aircraft model with two kinds of elastic wings and a rigid wing. The 3D deformations and 2D3C flow fields are measured using a synchronous measurement system, which includes the high accuracy deformation measurement (HADM) method and the stereoscopic particle image velocimetry (SPIV). It is shown that increased wing elasticity leads to greater bending, with the wing tip moving along the streamwise, vertically upward, and spanwise inward directions. These deformations are closely linked to aerodynamic forces and influence the tip vortex positions. The more elastic wing, EW2, exhibits the highest vortex core position, more than 0.5 times the length of the wing root, while the other one, EW1, shows the strongest vortex in the lift-increased regime. In the early wakes, the tip vortex of the elastic wing is elongated along the spanwise direction, resulting in an asymmetric shape. As it convects downstream, the vortex core disperses via viscous diffusion, resulting in a relatively more circular shape with decreased vorticity.

A Preliminary Evaluation of Morphing Horizontal Tail Design for UAVs

Morphing structures are a relatively new aircraft technology currently being investigated for a variety of applications, from civil to military. Despite the lack of literature maturity and its complexity, morphing wings offer significant aerodynamic benefits over a wide range of flight conditions, enabling reduced aircraft fuel consumption and airframe noise, longer range and higher efficiency. The aim of this study is to investigate the impact of morphing horizontal tail design on aircraft performance and flight mechanics. This study is conducted on a 1:5 scale model of a Preceptor N-3 Pup at its trim condition, of which the longitudinal dynamics is implemented in MATLAB release 2022. Starting from the original horizontal tail airfoil NACA 0012 with the elevator deflected at the trim value, this is modified by using the X-Foil tool to obtain a smooth morphing airfoil trailing edge shape with the same CLα. By comparing both configurations and their influence on the whole aircraft, the resulting improvements are evaluated in terms of stability in the short-period mode, reduction in the parasitic drag coefficient CD0, and increased endurance at various altitudes.

On the Aeroelasticity of a Cantilever Wing Equipped with the Spanwise Morphing Trailing Edge Concept

This paper studies the aeroelastic behavior of a rectangular, cantilever wing equipped with the spanwise morphing trailing edge (SMTE) concept. The SMTE consists of multiple trailing edge flaps that allow controlling the spanwise camber distribution of a wing. The flaps are attached at the wing’s trailing edge using torsional springs. The Rayleigh–Ritz method is used to develop the equations of motion of the wing-flap system. The use of shape functions allows for representing the wing as an equivalent 2D airfoil with generalized coordinates that are defined at the wingtip. Strip theory, based on Theodorsen’s unsteady aerodynamic model, is used to compute the aerodynamic loads acting on the wing. A representative Padé approximation for Theodorsen’s function is utilized to model the aerodynamic behaviors in a state-space form allowing time-domain simulation and analysis. The model is validated using a rectangular cantilever wing and the data are available in the literature. A comprehensive parametric comparison study is conducted to assess the impact of flap stiffness on the aeroelastic boundary. In addition, the potential of the SMTE to provide load alleviation and flutter suppression is assessed for a wide range of flight conditions, using a discrete (1-cosine) gust. Finally, the implementation and validation of a controller for a wing with SMTE for gust load alleviation are studied and controller parameters are tuned for a specific gust model.

Optimization Design of the NUAA-PTRE: A New Pre-Cooled Turbine Engine Adapting to 0~5 Mach Number

A model of a NUAA-PTRE pre-cooled air turbine engine was established. The design point parameters of the engine were optimized, including the pressure ratio, air flow rate of the compressor, efficiency, throat area, and efficiency of the turbine. The air flow rate at the engine operating point was 142.73 kg/s. High performance of the key components under a wide range of working conditions was realized after optimization. To achieve the indicators of the overall scheme, adaptability studies of key components were conducted. A three-stage variable geometry design was applied to the inlet. The pre-cooler was optimized with a power-to-weight ratio of over 100 kW/kg and a compactness of 278 m2/m3. The built-in rocket gas generator and dual-component injector were developed, and the combustion and heat transfer processes were simulated. The overall optimization design of the NUAA-PTRE and the adaptive design of the components were completed, and high performance of the engine in a wide range of flight conditions at Ma 0~5 and altitude 0~25 km was achieved.

Numerical study on initiation of oblique detonation wave by hot jet

Reliable detonation initiation is critical in detonation engines, especially for oblique detonation engines covering a wide range of flight regimes. The widely-studied oblique shock initiation usually leads to a significant variation in initiation length, making it difficult to adapt to the wide range of flight conditions. In this study, hot jets in the induction zone, i.e. behind the oblique shock, are introduced to explore the possibility of jet initiation, providing a potential method for initiation control. Results show that hot jet can decrease the initiation length to achieve rapid initiation and change the wave structure simultaneously, which is attributed to the oblique shock induced by a gaseous wedge. Thermal analyses of jet-initiated ODWs are also conducted through total pressure recovery. Results indicate that jet-initiated ODWs have a slightly better total pressure recovery than the original ODW. Furthermore, the jet angle and position are varied to analyze their effects on the initiation position, which is insensitive to the former while sensitive to the latter. A large jet angle may lead to a strong solution of secondary oblique detonation, which has not been previously reported.

Multiaxial mechanical characterization of latex skin for morphing wing application

A morphing aircraft continuously adjusts its wing geometry to achieve optimal flight characteristics for a wide range of flight conditions. Research on morphing aircraft has focused on developing wings with 1 morphing degree of freedom “monomorphing”. However, nowadays the focus is on developing wings with multiple morphing degrees of freedom “polymorphing”. One of the key design challenges for morphing aircraft technology is the development of skin capable of facilitating morphing whilst maintaining the aerodynamic shape of the wing. Soft polymeric materials represent a potential candidate to act as morphing skin because of their ability to undergo large deformation in multiaxial directions. In this study, latex skin is characterized under all possible deformation modes such as uniaxial, pure shear, biaxial, and equibiaxial to account for monomorphing and polymorphing applications. The effects of strain rate, thickness, and aspect ratio on hysteresis loss, stress relaxation and, stiffness are also studied. The outcomes of this study provide a comprehensive understanding of the mechanical viscoelastic behavior of latex skin under various deformation modes.

Morphology of oblique detonation waves in a stoichiometric hydrogen–air mixture

Abstract

Phenomenology and scaling of optimal flapping wing kinematics

Biological flapping wing fliers operate efficiently and robustly in a wide range of flight conditions and are a great source of inspiration to engineers. The unsteady aerodynamics of flapping wing flight are dominated by large-scale vortical structures that augment the aerodynamic performance but are sensitive to minor changes in the wing actuation. We experimentally optimise the pitch angle kinematics of a flapping wing system in hover to maximise the stroke average lift and hovering efficiency with the help of an evolutionary algorithm and in situ force and torque measurements at the wing root. Additional flow field measurements are conducted to link the vortical flow structures to the aerodynamic performance for the Pareto-optimal kinematics. The optimised pitch angle profiles yielding maximum stroke-average lift coefficients have trapezoidal shapes and high average angles of attack. These kinematics create strong leading-edge vortices early in the cycle which enhance the force production on the wing. The most efficient pitch angle kinematics resemble sinusoidal evolutions and have lower average angles of attack. The leading-edge vortex grows slower and stays close-bound to the wing throughout the majority of the stroke-cycle. This requires less aerodynamic power and increases the hovering efficiency by 93% but sacrifices 43% of the maximum lift in the process. In all cases, a leading-edge vortex is fed by vorticity through the leading edge shear layer which makes the shear layer velocity a good indicator for the growth of the vortex and its impact on the aerodynamic forces. We estimate the shear layer velocity at the leading edge solely from the input kinematics and use it to scale the average and the time-resolved evolution of the circulation and the aerodynamic forces. The experimental data agree well with the shear layer velocity prediction, making it a promising metric to quantify and predict the aerodynamic performance of the flapping wing hovering motion.

A neural network based MRAC scheme with application to an autonomous nonlinear rotorcraft in the presence of input saturation

This paper develops a neural-network-based model reference adaptive control (MRAC) scheme for a rotorcraft in the presence of input saturation. Such a control scheme provides acceptable tracking performance for the rotorcraft in a wide range of flight conditions. Combined with hyperbolic tangent functions, the MRAC scheme is capable of tracking the reference signals without violating input constraints. A modified projection operator is utilized to prevent system from parameter drift due to the strong nonlinearity and uncertainty of the rotorcraft mathematical model. Stability of the proposed MRAC scheme is proved based on Lyapunov stability theory. The performance of the resulting controller is tested by conducting numerical simulations for an autonomous rotorcraft in various flight missions.

Nonlinear sliding sector design for multi‐input systems with application to helicopter control

SummaryThe ability of helicopters to hover and land vertically has spurred an interesting field of research on the development of autonomous flight for these rotatory wing aircrafts. Linear control theory with gain scheduling, which is based on linearizing the system at the equilibrium points, dominated the helicopter autopilot design. Unlike the linear cascaded autopilot structure used in the existing literature, this paper uses state‐dependent linear like structure, including rate‐limited actuator dynamics, with cascaded autopilot topology. This approach allows nonlinear control laws to be implemented throughout the entire flight envelope, providing satisfactory robustness and stability over the various parameter uncertainties and time delays. The cascaded autopilot topology with nonlinear dynamical equations contains a new sliding sector control (SSC) mechanism which is derived for multi‐input nonlinear dynamical systems. The proposed SSC structure for multi‐input nonlinear systems is used in the inner loop of the cascaded autopilot system where the fastest dynamics are required to be controlled for rapid changes in the helicopter dynamical characteristics which enables one to stabilize the helicopter over a wide range of flight conditions. The proposed cascaded autopilot topology with the new SSC mechanism is tested in simulations to assess its robustness and stability properties. To establish its feasibility, the proposed control method is replaced with a suboptimal control method, namely state‐dependent differential Riccati equation (SDDRE) method, for the inner loop and the results of the proposed control architecture are compared with those of SDDRE method.

Control device effectiveness studies of a 53∘ swept flying wing configuration. Experimental, computational, and modeling considerations

The present investigation covers studies of different control devices on several 53∘ swept flying wing configurations without a vertical tail plane. The wings considered share the same planform but differ in their spanwise profile shapes. The planform, developed under the NATO AVT-251 Task Group, is referred to as the MULDICON (or MULti-DIsciplinary CONfiguration). The objectives of this article are twofold: (1) to design yaw control surfaces for the MULDICON wing using an experimental/computational approach and (2) to develop aerodynamic models that rapidly and accurately predict the effectiveness of control surfaces over a wide range of flight conditions. The yaw control surface design (position and size) should provide sufficient yaw moment with almost no contribution in roll and pitch moment. To identify such concepts, a number of preliminary experiments on a generic flying wing configuration have been conducted. Two promising concepts from wind tunnel tests were then numerically examined for being implemented in the MULDICON baseline wing. Concepts with spoilers and a split flap were specifically considered. For medium to high angles of attack, the flow topology of the baseline wing is dominated by a vortical flow field on the upper outer wing. This leads to interactions between vortex and the control device which influences the flow and the attitude of the control device on the upper wing side. Furthermore, this work considers developing aerodynamic models for predicting stability derivatives of several MULDICON designs over a wide range of flight conditions. Aerodynamic loads models are only developed for normal force and pitch moment coefficients, however, the developed approach can easily being extended to include lateral aerodynamic coefficients as well. Quasi-steady models of this work are power series expansions of traditional linear aerodynamic models to capture nonlinear effects. Additionally, the models can estimate static and dynamic stability derivatives and the control surface powers.

The potential of helicopter turboshaft engines incorporating highly effective recuperators under various flight conditions

Incorporating recuperators into gas turbines shows considerable potential for lower emissions and fuel consumption. Nowadays the technology readiness of advanced compact heat exchanger has provided a solid foundation for the availability of lightweight, higher efficient recuperators which would find good acceptance on the rotorcraft without penalizing the operational capabilities. To understand the impact of recuperator on the whole system for further development of future recuperated helicopter, it is proposed to evaluate the potential of recuperated helicopter turboshaft engines with emphasis placed on highly effective primary surface recuperator. This paper presents a comprehensive multidisciplinary simulation framework, and the aircraft configuration selected is a generic helicopter, which is similar to the helicopter Bo105, equipped with two Allison 250-C20B turboshaft engine variants. The improved part-load performance against the reference non-recuperated cycle is discussed first, followed by the analysis and evaluation of two representative flight missions. The study is finally extended to quantify the flight time required to compensate for the additional recuperator weight under the flight condition of 0-250 km/h and 0-3000 m for different recuperator design effectiveness values. It is suggested that the selection of recuperator effectiveness should be dependent on the most commonly involved mission profile and flight duration, in order to offset the added parasitic weight of the recuperator. The established rotorcraft multidisciplinary framework proves to be an effective tool to conduct a comprehensive assessment for the recuperated helicopter under a wide range of flight conditions as well as at mission level.

The Stability of the Optimal Aerodynamic Design of an Isolated Three-Dimensional Wing to Its Initial Form

The results are presented for the initial form stability analysis of the optimal aerodynamic design algorithm in context of an isolated three-dimensional wing of a wide-body long-range aircraft. The solution to the problem of geometry determination with a minimum of wing total drag subject to a fixed lift coefficient with allowance for numerous aerodynamic and geometrical constraints by means of the algorithm combining high-precision mathematical modeling and global optimal search using supercomputing technologies. It is established that the algorithm is stable to the choice of the wing initial form, because the optimal designs obtained for two considerably distinct variants of the wing’s initial form are very close to each other and have practically identical integral aerodynamic properties at the main design point as well as in a wide range of flight conditions.

Investigation of Smart Material Actuators & Aerodynamic optimization of Morphing Wing

The design of wings in traditional aircraft creates a configuration with optimal performance but that suits to a single flight condition only. As the vehicle moves away from the given condition performance too declines thereof. Conversely, birds can adopt their wing profile to augment the performance at a wide range of flight conditions. Morphing technology offers aerodynamic proficiency over a wide range of flight conditions, and the key to enable this technology for adaptive structures is based on smart materials and actuators. This smart technology has impended to the development of multi-mission aerospace vehicles by shaping the adaptive wings that can adjust with the operating maneuvers. In this work, the analyses of adaptive wing section are performed by attaining a variable span geometry to increase the efficiency of an aircraft. The NACA4412 aerofoil is introduced to CFD solver, FLUENT. The aerodynamic coefficients are obtained at various angles of attack in both the extended (morphed) wing section as well as un-extended wing profile (un-morphed). Results show that the extended wing section increases the aerodynamic performance of wing profile by varying its aspect ratio at different angles of attack.

Gas-dynamic problems in off-design operation of supersonic inlets (review)

Modern concepts of operation of supersonic inlets of high-velocity air-breathing engines are analyzed. It is demonstrated that the flow in the engine duct becomes extremely complicated in off-design modes of inlet operation, which can lead to unpredictable consequences, in particular, to inlet unstart. The term “inlet unstart” is considered in the present paper as a synonym of the absence of theoretical understanding and prediction of gas-dynamic phenomena. Various approaches are proposed to ensure self-regulation of the inlet-combustor system for air-breathing engines. Possible directions of further research are indicated for the purpose of stable operation of inlets in a wide range of flight conditions.

F-16XL Simulations at Flight Conditions Using Hybrid Near-Body/Offbody Computational Fluid Dynamics

The F-16XL flight vehicle has been used extensively to study the aerodynamics of medium- to high-angle-of-attack flight. Flight-test data, including surface pressures, are available for a wide range of flight conditions, allowing direct comparison between computational fluid dynamics and full-scale flight data. The most recent comprehensive study, made possible by NASA, was conducted by NATO Task Group AVT-113 and had participants from many different countries. One of the conclusions of the study was that unsteady flow simulations with high-resolution turbulence treatment was necessary to compare well with the high-angle-of-attack flight-test data. The current work applies the high performance computing CREATE™-Air Vehicles Kestrel fixed-wing simulation tool to the F-16XL configuration at high-angle-of-attack flight conditions. Kestrel couples a near-body spatially second-order solver with an offbody, spatially third- or fifth-order solver to achieve a very high-resolution computation of the aerodynamic features of the flowfield. Adaptive mesh refinement is also performed in the offbody domain as the solution progresses to maintain the vortical structures away from the body. The resulting simulation data are compared to flight-test data, and they are found to be in very good agreement for this flight condition.

Morphing aircraft based on smart materials and structures: A state-of-the-art review

A traditional aircraft is optimized for only one or two flight conditions, not for the entire flight envelope. In contrast, the wings of a bird can be reshaped to provide optimal performance at all flight conditions. Any change in an aircraft’s configuration, in particular the wings, affects the aerodynamic performance, and optimal configurations can be obtained for each flight condition. Morphing technologies offer aerodynamic benefits for an aircraft over a wide range of flight conditions. The advantages of a morphing aircraft are based on an assumption that the additional weight of the morphing components is acceptable. Traditional mechanical and hydraulic systems are not considered good choices for morphing aircraft. “Smart” materials and structures have the advantages of high energy density, ease of control, variable stiffness, and the ability to tolerate large amounts of strain. These characteristics offer researchers and designers new possibilities for designing morphing aircraft. In this article, recent developments in the application of smart materials and structures to morphing aircraft are reviewed. Specifically, four categories of applications are discussed: actuators, sensors, controllers, and structures.

Quasi-steady aerodynamic model of clap-and-fling flapping MAV and validation using free-flight data

Flapping-wing aerodynamic models that are accurate, computationally efficient and physically meaningful, are challenging to obtain. Such models are essential to design flapping-wing micro air vehicles and to develop advanced controllers enhancing the autonomy of such vehicles. In this work, a phenomenological model is developed for the time-resolved aerodynamic forces on clap-and-fling ornithopters. The model is based on quasi-steady theory and accounts for inertial, circulatory, added mass and viscous forces. It extends existing quasi-steady approaches by: including a fling circulation factor to account for unsteady wing–wing interaction, considering real platform-specific wing kinematics and different flight regimes. The model parameters are estimated from wind tunnel measurements conducted on a real test platform. Comparison to wind tunnel data shows that the model predicts the lift forces on the test platform accurately, and accounts for wing–wing interaction effectively. Additionally, validation tests with real free-flight data show that lift forces can be predicted with considerable accuracy in different flight regimes. The complete parameter-varying model represents a wide range of flight conditions, is computationally simple, physically meaningful and requires few measurements. It is therefore potentially useful for both control design and preliminary conceptual studies for developing new platforms.

Aircraft suitability model applied to a laminar flow control flight experiment

When planning a flight experiment, the test team must analyse the suitability of the testbed aircraft. Often, critical design decisions for the test apparatus are made without high-fidelity models of aircraft performance since few testbed aircraft are designed and built for the flight experiment. This study develops a process to adapt rough performance data to determine the suitability of an airplane to operate at a given angle of attack across a wide range of flight conditions within operational limits. The process illustrates the careful balance of factors in the design of a swept wing laminar flow control wing glove. The process also establishes an important flight planning tool that allows the test team to assess the impact of aircraft performance on the flight experiment. Further, the tool can be updated with flight data when available and leveraged as a planning tool to promote safe and efficient flight research.