Modes Of Aircraft Research Articles

THE KEY ROLE OF MODERN AERODYNAMIC TRENDS IN INCREASING THE ENERGY EFFICIENCY OF HIGH-SPEED VEHICLES

Based on the analysis of typical operating modes of modern high-speed trains and transport aircraft, the influence of various schemes of surface relief modification and blowing through it on aerodynamic drag has been thoroughly investigated. A set of several schemes of the combined effect of these two factors is proposed to create an easy-to-use, cheap-to-produce streamlined surface of the desired structure with a reliable and controllable effect of active influence on aerodynamic drag reduction. The potential effect of reducing aerodynamic drag has been verified by numerical and Particle Image Velocimetry experimental modeling. The proposed technological innovations are encouraging and relevant; by reducing energy costs, they increase the commercial and environmental attractiveness of their implementation.

METHOD OF ESTIMATING THE DANGEROUS RANGE OF TRANSONIC NUMBERS M OF THE FLIGHT OF SUPERSONIC AIRCRAFT AND AEROSPACE SYSTEMS

Ensuring flight safety of supersonic aircraft and aerospace systems in the transonic range of flight M numbers still remains an ac- tual scientific and applied problem. This is due to the occurrence of various dangerous aeroelasticity phenomena in these flight modes. Such phenomena include transonic flutter, the occurrence of which has repeatedly resulted in the destruction of aircraft structural elements and, first of all, of aerodynamic control surface structural elements. Many publications are devoted to the theoretical and experimental research of this phenomenon, in which various approaches are proposed to substantiate the causes of intense oscillations of the aerodynamic control surfaces of modern supersonic aircraft in these flight modes, the conditions of their occurrence, the influence of various factors on the level of oscillations. It is noted that there are still no reliable theoretical methods for estimating the non-stationary forces of aerodynamic control surfaces that oscillate in a transonic flow, so the use of linear mathematical similarity models does not always allow transferring the results of blowing models in wind tunnels to full-scale aircraft designs. The paper proposes a method for estimating the dangerous range of M numbers in which transonic flutter of the aerodynamic control surfaces of supersonic aircraft and aerospace systems is possible. The method is based on the analysis of regularities of the adiabatic expansion of the local supersonic air flow on the surface of the airfoil in the range of transonic numbers M. Calculations have proven that for typical aerodynamic surfaces of modern supersonic aircraft, the occurrence of transonic flutter is possible in a narrow range of numbers M = 0.9…0.94. The obtained results can be used to substantiate the safe flight modes of supersonic aircraft both at the stage of flight tests and at the stage of operation. Further studies of this problem should be devoted to the analysis of methods for reducing the level of oscillations of aerodynamic control surfaces when transonic flutter occurs.

On ecologically-safe high-speed vehicles: Conceptual design study of the next generation supersonic transport

Space-age technology makes it possible to travel faster than the speed of sound from one continent to another one or even to make suborbital flight as a space tourism. The existing technology challenge is to bring this innovative way of travel without damaging the environment to ensure human safety and to preserve environment balance. The sonic boom issue is one of the concerns when flying faster than sound. The issue must be taking into account at the stage of developing environmentally friendly high-speed vehicle. The next generation supersonic transport is fundamentally different from subsonic medium- and long-range aircraft in a breakthrough feature: it ensures a long-term supersonic cruise flight in the range of Mach numbers 1.6–1.8. For a successful implementation of the civil supersonic project at a new stage of technology development, it is necessary to solve a number of scientific and technical problems. In the present paper some of them are under discussion. In particular, the addressed issues are as follows: how to achieve a high level of aerodynamic efficiency at supersonic cruise along with low sonic boom; how to fit into low noise requirements at take-off and landing modes of supersonic aircraft with a moderate-bypass-ratio engine; how to implement a wing with high load factor and fuselage with long nose part to fit into required airframe structural stiffness, etc. Also, new technology approaches in multidisciplinary optimization task to achieve aircraft design goals of low drag, low boom and low noise characteristics are discussed. A few examples of optimal supersonic aircraft configurations along with passenger cabin are given. The paper was presented at the X-th IAA Symposium “Space Flight Safety” in St.-Petersburg on June 2023.

Experimental Study on Modal Characteristics of Large Flexible Aircraft

High altitude long endurance unmanned aerial vehicle (UAV) is a type of UAV that operates at an altitude of around 10000 meters and can fly for several days, relying on solar panels.Usually, this type of aircraft has the structural characteristics of high aspect ratio, large wingspan, low wing load, and high flexibility.In engineering applications, it is often necessary to measure the vibration modes of such aircraft, and the modal measurements of such aircraft are often not accurately obtained.This article conducts relevant research on the issue of inaccurate measurement in modal experiments of such aircraft. By establishing a finite element model, the reasons for the inaccurate measurement of low order frequencies in modal experiments of such aircraft are analyzed, and the magnitude of data errors in different situations is quantitatively calculated.Based on the above research, this article proposes two methods for solving modal tests of such aircraft. This method has important guiding significance for ground modal testing of large flexible structures, which is beneficial for dynamic analysis and flight control research of such aircraft.

Determination of the most dangerous flight modes of aircraft in icing conditions

In this paper, the object of research is the icing of aircraft surfaces during flight in the atmosphere. On many light aircraft, as well as on unmanned aircraft weighing less than 30 kg, there are no on-board de-icing systems. Nevertheless, aviation events occur with these aircraft, which are a consequence of their icing. Therefore, determining the most dangerous flight modes of aircraft in icing conditions is an urgent task. In view of the high cost of conducting flight tests and the impossibility of covering all possible events due to their potential danger, the complexity of creating flight conditions for aircraft in icing conditions on the ground, the mathematical modeling method was used in this study. To solve this problem, the analysis of the airworthiness standards of civil light aircraft, transport category aircraft, rotorcraft of normal and transport category was carried out within the framework of the work, the influence of various parameters on the thickness of ice build-up was investigated using a computational experiment conducted on the software developed by the authors of the article. On the basis of the results of the computational experiment, the dependences of the ice thickness on various icing parameters were obtained, a method was developed for determining the combination of heights and flight speeds of an aircraft, at which ice of the greatest thickness is formed on the surface of aircraft, other things being equal. Possession of this information will allow the aircraft crew and air traffic control specialists to avoid the most dangerous flight modes in terms of icing.

Modal Analysis and Flight Validation of Compound Multi-Body Aircraft

The flight dynamics of a compound aircraft are highly coupled and significantly different from those of a single aircraft. In this paper, the characteristic flight dynamics of a three-unit connected aircraft are analyzed in detail. By distinguishing the aircraft’s modes into a symmetric mode group and an asymmetric mode group, the mechanisms by which the relative roll modes of the compound aircraft significantly impact the rigid-body modes of the vehicle become clear. The trim configuration of a compound aircraft has a profound influence on its stability and flight modes. In addition, the relative roll modes can be separated from the full-state model by appropriate simplifications, and the modal coupling mechanism and evolution law can be analytically derived. The key parameters which have a significant influence on the trim configuration and flight modes of the aircraft are pointed out in this paper. Finally, the common failure modes of compound aircraft are identified. A flight experiment with test data verifies the correctness of the above theoretical analysis.

Co-design Optimization of a Novel Multi-identity Drone Helicopter (MICOPTER).

Delivery drones have always faced challenges when it comes to reliably deliver packages. This paper introduces a novel concept of a hybrid drone called “MICOPTER” to alleviate this issue. Being able to fly in three modes of aircraft, helicopter, and gyrocopter, the proposed model of the multi-identity helicopter comprises a 2DOF tilting mechanism of rotors and a folding wing system leading to better performance and controllability. To scrutinize the idea, MICOPTER is compared to other types of Unmanned Aerial Vehicles (UAVs) in terms of different performance parameters. The performance goal of the MICOPTER is the realization of a predetermined standard delivery drone mission based on Amazon Prime Air. According to the relevant literature, the corresponding conceptual design equations are formulated and the traditional matching diagram method is utilized to attain the initial design point. Afterward, a multidisciplinary-feasible design matrix is provided as well as multi-objective optimization to strive for optimal feasible configurations while maximizing cruise velocity and range. Furthermore, the configuration and performance of some of the feasible design points on the final Pareto frontier are compared with the traditional design. Finally, by simulating a typical flight profile and using robust non-linear backstepping control, the controllability of the proposed configuration is investigated. The controller performance is assessed considering its stability and tracking 8-shape trajectory. Results indicate the MICOPTER capabilities as a novel configuration in both terms of design performance and controllability.

Objective comparison of numerical spin study with aircraft model free-flight spin tests

This paper reports on the results of a research project that investigated the spin and recovery characteristics of a multirole fighter aircraft. In the first phase, numerical/computational methods were used to predict full scale aircraft spin and recovery characteristics. The aerodynamic data used in the numerical study was obtained from rotary balance steady coning and oscillatory coning motion dynamic wind tunnel tests conducted on 1/13th scale aircraft dynamic model. All the possible steady spin modes were numerically computed, a complete set of dynamic modes of the spinning aircraft were evaluated and six degree of freedom simulations of predicted flat spins were performed to investigate their dynamic stability and recovery characteristics. In the second phase, free-flight spin trials were conducted using 1/5th scale dynamic model of the understudy aircraft to validate the results of the analytical spin studies. Dynamic characteristics of the analytically predicted flat spins were objectively compared with flat spins found in free-flight spin trials as a validation process. Our results reveal that the numerically predicted full scale aircraft spin and recovery dynamic characteristics correlate well with those estimated from model test flights. The present study is an important step toward objective comparison of numerical spin studies based on dynamic wind tunnel tests data with free-flight spin trials, yielding important guidelines and results for future aircraft spin studies.

TO THE PROBLEM OF MODELING THE AERODYNAMICS OF SURFACE TRANSPORTATION VEHICLES

The article deals with the problems of modeling the aerodynamics of surface high-speed vehicles of winged layout moving at a short distance from the water surface. To support them in the air, the screen effect formed under the influence of proximity to the water surface should be used. The effect of ground proximity is observed in air transport in the takeoff and landing modes of aircraft. The proximity of the ground affects the formation of the aerodynamic characteristics of vehicles. Today, the study of the aerodynamics of advanced vehicles moving near the underlying surface is possible using the following approaches: full-scale tests; modeling on special tracks; wind tunnel modeling. Full-scale tests are usually used at the final stages of vehicle development. simulation on special tracks is associated with the complexity of obtaining aerodynamic characteristics and is not widely used. The most suitable is modeling using wind tunnels. The current level of equipment of experimental laboratories allows for correct aerodynamic studies of aircraft. However, when modeling aerodynamic processes around vehicles moving near the interface, complex problems arise with reproducing the conditions of motion dynamics. The paper describes in detail the methods of studying the influence of the proximity of the ground on the formation of aerodynamic characteristics of vehicles in wind tunnels. These include: a method using a fixed plate; a method of mirroring the model; a method of suction or blowing the boundary layer; a method using a moving belt. The possibility of their use for modeling the aerodynamics of high-speed surface vehicles of winged layout is evaluated. When modeling the flow of high-speed vehicles moving near the water surface, a number of difficulties arise due to the realization of real flow conditions, namely, taking into account the dynamics of the water surface. Ideally, relative motion is realized between the vehicle and the water surface. The review of methods used in modeling the aerodynamics of vehicles over a solid screen proved to be of little use for high-speed winged vehicles moving at a short distance from the water surface. Ensuring the similarity of relative motion is an extremely difficult task. It requires a reasonable analysis of physical processes and the search for ways to develop physical and mathematical models suitable for use in the design of high-speed surface vehicles with wing arrangement. These models should range from simple engineering techniques to techniques based on the full and averaged Navier-Stokes equations.

Design of sliding mode flight control system for a flexible aircraft

The evolution of large transport aircraft is characterized by longer fuselages and larger wingspans, while efforts to decrease the structural weight reduce the structural stiffness. Both effects lead to more flexible aircraft structures with significant aeroelastic coupling between flight mechanics and structural dynamics, especially at high speed, high altitude cruise. The lesser frequency separation between rigid body and flexible modes of flexible aircraft results in a stronger interaction between the flight control system and its structural modes, with higher flexibility effects on aircraft dynamics. Therefore, the design of a flight control law based on the assumption that the aircraft dynamics are rigid is no longer valid for the flexible aircraft. This paper focuses on the design of a flight control system for flexible aircraft described in terms of a rigid body mode and four flexible body modes and whose parameters are assumed to be varying. In this paper, a conditional integral based sliding mode control (SMC) is used for robust tracking control of the pitch angle of the flexible aircraft. The performance of the proposed nonlinear flight control system has been shown through the numerical simulations of the flexible aircraft. Good transient and steady-state performance of a control system are also ensured without suffering from the drawback of control chattering in SMC.

Advanced Navigation Control Systems

Compared to the other major modes of transportation aircraft is the fastest mode till now with few challenges. The major and important challenge is the safety of passengers. A small mistake regarding safety operations in aviation leads to demise of passengers and crew. In recent times and in history many flights had gone missing and highjacked. Even the information about some of the flights were unknown till now. We even had a history about the airlines that pay-off a huge indemnity for an air crash. Various navigation systems are present with the advanced navigation techniques from GPS to satellite navigation with many limitations. There are even many safety considerations and advanced techniques that are in effect for the safety of passengers but the plane crashes had not been nullified. There is a need to improve the safety procedures and navigation systems for the safety of flights. This paper presents the information about the safe navigation systems that enables the safe flight. These methods and techniques are suggested based on the research and analysis of various missing flights and flight crashes. Using these navigation systems, highjack, flight crashes, missing flights can be reduced to a great extent. By incorporating these methods and systems on board, flight will be safer.

"THEORETICAL ASPECTS OF STRENGTH AND THERMAL CONTROL OF HYPERSONIC AIRCRAFT"

This article analyzes the theoretical aspects of controlling the strength and thermal modes of hypersonic aircraft. The conditions for the functioning of hypersonic aircraft were also investigated, and problematic situations for their design were identified. 
 The parameters of aerodynamic heating of surfaces and heating of thermal protection of hypersonic aircraft were estimated with an assessment of the parameters of thermal protection of hypersonic aircraft and the magnitude of the heat flux supplied to the surface, with the determination of the parameters of their thermal protection, taking into account the thermophysical characteristics of materials from thermal parameters.

Study on Control Strategy for Tilt-rotor Aircraft Conversion Procedure

This paper studies the control strategy in tilt-rotor aircraft dynamic conversion procedure. A nonlinear flight dynamics model with full flight modes is established. On this basis, a nonlinear optimal control model of dynamic conversion is built by constructing the Bolza form of nonlinear optimal control problem. It contains the limitations and effects of conversion corridor, pilot control, flight attitude, engine rated power, wing stall, and the cooperation between lift and thrust on the procedure of dynamic conversion. An efficient numerical solution method with good convergence is designed to obtain the trajectory and control strategy of dynamic conversion between the modes of helicopter and fixed-wing aircraft. Through the weight parameter analysis, it can be seen that when the subitem of pilot workload is included in the cost function, the displacements of collective stick input as well as the longitudinal stick input are significantly reduced, and the height and pitch attitude change more gently, but the time has been extended. In addition, the pilot workload weight coefficient should not be dominant, otherwise the overly smooth manipulations of collective and longitudinal stick will cause a decrease of height.

Formation Mission Design for Commercial Aircraft Using Switched Optimal Control Techniques

In this article, the formation mission design problem for commercial aircraft is studied. Given the departure times and the departure and arrival locations of several commercial flights, the relevant weather forecast, and the expected fuel savings during formation flight, the problem consists in establishing how to organize them in formation or solo flights and in finding the trajectories that minimize the overall direct operating cost of the flights. Each aircraft can fly solo or in different positions inside a formation. Therefore, the mission is modeled as a switched dynamical system, in which the discrete state describes the combination of flight modes of the individual aircraft and logical constraints in disjunctive form establish the switching logic among the discrete states of the system. The formation mission design problem has been formulated as an optimal control problem of a switched dynamical system and solved using an embedding approach, which allows switching decision among discrete states to be modeled without relying on binary variables. The resulting problem is a classical optimal control problem which has been solved using a knotting pseudospectral method. Several numerical experiments have been conducted to demonstrate the effectiveness of this approach. The obtained results show that formation flight has great potential to reduce fuel consumption and emissions.

An improved online system identification method for tiltrotor aircraft

An online system identification method for tiltrotor aircraft flight dynamics modeling is developed in this paper. A general parametric flight dynamics model which can be used for online identification purpose in the three flight modes of tiltrotor aircraft (i.e., helicopter mode, conversion mode, and fixed-wing mode) is developed, and an unideal noise model is also introduced in order to minimize the parameter identification error caused by measurement noise. An adaptive model structure identification algorithm is established by introducing local accuracy criteria, so that a time variant optimal model structure is obtained in real time. A weighted recursive least squares algorithm is established for parameter identification, and a time variant weighting matrix is designed to provide best numerical performance in the identification procedure at any time. A comprehensive nonlinear flight dynamics model of a notional tiltrotor aircraft is used to generate simulation flight test data in all flight modes, and the identification results of the notional tiltrotor aircraft in fixed wing mode, helicopter model as well as conversion mode are obtained and verified. Finally, the flight dynamics model of an unmanned tiltrotor aircraft in helicopter mode is identified by using the real flight test data. The identification results show that the online system identification method for tiltrotor aircraft developed in this paper is effective, efficient and robust.

Aeroelastic scaling of flying demonstrator: flutter matching

The traditional approach for the design of aeroelastically scaled models assumes that either there exists flow similarity between the full-size aircraft and the model, or that flow non-similarities have a negligible effect. However, when trying to reproduce the behavior of an airliner that flies at transonic conditions using a scaled model that flies at nearly-incompressible flow conditions, this assumption is no longer valid and both flutter speed and static aerodynamic loading are subject to large discrepancies. To address this issue, we present an optimization-based approach for wing planform design that matches the scaled flutter speeds and modes of the reference aircraft when the Mach number cannot be matched. This is achieved by minimizing the squared error between the full-size and scaled aerodynamic models. This method is validated using the Common Research Model wing at the reference aircraft Mach number. The error in flutter speed is computed using the same wing at model conditions, which are in the nearly-incompressible regime. Starting from the baseline wing, its planform is optimized to match the reference response despite different conditions, achieving a reduction of the error in the predicted flutter speed from 7.79% to 2.13%.

БАЗОВОЕ ИСПОЛНЕНИЕ РЕГУЛЯТОРА АВИАДВИГАТЕЛЯ И ЕГО МОДИФИКАЦИИ

It is shown that the creation of a unified electronic control and control equipment for aircraft gas turbine engines is one of the determining trends in the experimental design work of enterprises developing components of aircraft equipment. The results of the creation of some helicopter and aircraft engine controllers (turboshaft and turboprops) in JSC Element are presented. The basic model for them was the regulator RDC-450М (category A component product according to AP-21 classification), which was developed for the AI-450M family of turboshaft engines of Мi-2M and Mi-2MSB and in 2014 received the APPLIANCE DESIGN APPROVAL. There is given information concerning the mass-produced regulator RDTs-450M rework to meet the new requirements specified by the engine designer as the results of certification tests of the AI-450M engine and state tests of the Mi-2M and Mi-2MSB helicopters. It is shown that the rework was carried out taking into account the maximum possible “modifiability” of the regulator, which is the maximum ease of its adaptation in the future to new requirements, including requirements for working with other types of engines, including turboprops. The modular principle of constructing the functional structure and design of the regulator is described; its structural scheme is given. Data are given on the results of equivalent cyclic tests of the regulator RDTs-450M, which was modified to meet the new requirements, and on subsequent work on its adaptation to monitor and regulate the modes of turboprop aircraft engines AI-450C (including its modifications) and MS-500V-02S. Information is given on the newly developed regulators RDTs-450M-S, RDTs-450M-S-1, RDTs-450M-S-T-P, RDTs-450S-500, among which there is an engine regulator designed for an unmanned aerial vehicle. It is noted that the modular design assumes the presence of a significant number of board-to-board connections, and the prospects of increasing the reliability of regulator having a modular design by using the technology of flexible-rigid boards are described.

Reduced Order Model Based Flight Control System for a Flexible Aircraft

The interaction between the flight control system and its structural modes becomes stronger if the frequency separation between rigid body mode and flexible modes of flexible aircraft are closer, and it introduces higher flexibility effects on aircraft dynamics. In such a case, the design of a flight control law is not valid based on the assumption that aircraft dynamics are rigid body models. Moreover, an integrated aircraft model having a rigid body and elastic body modes contain a large number of states. Therefore, we have designed an optimal flight control law from a reduced-order model and realized with a flexible aircraft represented by a full order model. For this, the simplified model of the flexible aircraft is derived using the Chidambara technique, and Luenberger observer is applied to estimating the elastic states of a simplified model from the aircraft measurements.

Параметрическое исследование конвективного теплообмена при обтекании затупленного конуса сверхзвуковым воздушным потоком

The paper focuses on a mathematical model for the parametric calculation of convective heat transfer on the surface of a blunted cone with the supersonic airflow around it. We compared the computational data with the known experiments on the cone airflow with Mach number М∞ = 8, and found the parameters of the computational grid, which allow us to have less deviation in calculation results from the experimental data. Taking into account the recommendations developed for the selection of the computational grid, we carried out the parametric study of the supersonic airflow around cones with the different half - angels in the range of Mach numbers from 4 to 10. According to the findings of the study, we made the database that describes the influence of the parameters on the coefficient of heat transfer from the supersonic airflow to the cone surface. The data obtained can be used to optimize the geometric configuration and flight modes of high-speed aircraft.

The Longitudinal Dynamic Characteristics of a V/STOL Aircraft Based on Deflected Slipstream during the Hovering Phase

Compared with general V/STOL (vertical/short taking-off and landing) aircrafts, the V/STOL aircraft based on deflected slipstream shows extremely different dynamic characteristics during the hovering phase. In this paper, a certain new type of V/STOL aircraft based on slipstream deflection is designed as research object. According to a simplified longitudinal dynamic model for the V/STOL aircraft based on deflected slipstream, it is proved that trim flap angle and centre of gravity position are two important factors during the hovering phase. Different flap angle means different trim state while the variation of centre of gravity position could change the range of trimming. Besides, modes of aircraft during the hovering phase are drawn and the law of eigenvalues changing as the variation of trim flap angle and center of gravity position is discovered. Regarding the findings as reference, the design of V/STOL based on deflected slipstream and control law design for V/STOL based on deflected slipstream during the hovering phase can be greatly improved.