Large Civil Aircraft Nonlinear Flight Dynamics Iteration Trim Method Based on the Downhill Simplex Optimization Algorithm

PurposeThe purpose of this paper is to present a novel GREY‒ASMAA model for reliability growth evaluation in the large civil aircraft test flight phase.Design/methodology/approachAs limited data are collected during the large civil aircraft test flight phase, which are not enough to meet the requirements of the ASMAA model for reliability growth, four basic GM(1, 1) models, even grey model, original difference grey model, even difference grey model and discrete grey model, are presented. Then both forward and backward grey models GM(1,1) are built to forecast and obtain virtual test data on left and right sides. Then the ASMAA model for reliability growth evaluation can be built based on original and virtual test data.FindingsAiming at the background of poor information data during the large civil aircraft test flight phase, first, a novel GREY‒ASMAA model, which was combined by the grey model GM(1,1) with the ASMAA model, has been put forward in this paper.Practical implicationsThe GREY‒ASMAA model for reliability growth evaluation can be used to solve the problem of reliability growth evaluation with poor information data during the large civil aircraft test flight phase, and it has been used in reliability evaluation of C919 at the test flight stage.Originality/valueThis paper presents two new definitions of forward grey model GM(1,1) and backward grey model GM(1,1), as well as a novel GREY‒ASMAA model for reliability growth evaluation of large civil aircraft during test flight phase.

The bifurcation and continuation methodology has evolved over the last two decades into a powerful tool for the analysis of trim and stability problems in aircraft flight dynamics. Over the years, bifurcation methods have been employed to deal with a variety of problems in aircraft dynamics, such as predicting high angle of attack behavior, especially spin, and studying instabilities in rolling maneuvers. The bifurcation methodology has served as a tool for the design of flight control systems, and is promising to be a useful tool in the aircraft design, simulation, testing, and evaluation process. In the present paper, we describe the state-of-theart in the use of bifurcation and continuation methods for the analysis of aircraft trim and stability with a few illustrative examples. Both the standard and extended bifurcation analysis procedures are discussed and typical results for instabilities in high-α flight and inertia-coupled roll maneuvers are shown. This is followed by several problems in nonlinear flight dynamics where bifurcation and continuation methods have been fruitfully applied to yield effective solutions. Finally, the use of bifurcation theory to arrive at analytical instability criteria is demonstrated for the aircraft roll coupling and wing rock problems. 78 references have been cited in the text.

The bifurcation and continuation methodology has evolved over the last two decades into a powerful tool for the analysis of trim and stability problems in aircraft flight dynamics. Over the years, bifurcation methods have been employed to deal with a variety of problems in aircraft dynamics, such as predicting high angle of attack behavior, especially spin, and studying instabilities in rolling maneuvers. The bifurcation methodology has served as a tool for the design of flight control systems, and is promising to be a useful tool in the aircraft design, simulation, testing, and evaluation process. In the present paper, we describe the state-of-the-art in the use of bifurcation and continuation methods for the analysis of aircraft trim and stability with a few illustrative examples. Both the standard and extended bifurcation analysis procedures are discussed and typical results for instabilities in high-α flight and in inertia-coupled roll maneuvers are shown. This is followed by several problems in nonlinear flight dynamics where bifurcation and continuation methods have been fruitfully applied to yield effective solutions. Finally, the use of bifurcation theory to arrive at analytical instability criteria is demonstrated for the aircraft roll coupling and wing rock problems. 76 references have been cited in the text.

WTO Case Law in 2013

The flight dynamics of aircraft were derived from Newton’s Law of Mechanics and Euler’s Equation of Moment in kinematics. These nonlinear equations were relatively complex to be solved analytically. In the other hand, they were also costing high computational load to be solved in conventional numerical method. Due to the practical need to analyze the aircraft flight dynamics, some simplified method has emerged in solving the equations with tolerable margin of errors. Most of these simplified methods were developed earlier in conventional aircraft configuration, which consist of fuselage (body), wing, and clearly defined vertical tail and horizontal tail. Recently, the BPPT (The Agency of The Assessment and The Application of Technology, Indonesia) has prototyping several type of UAV. The development was including the unconventional configuration, such as the V-Tail UAV type named BPPT-01B “Gagak” (“Crow” in English). The V-tail configuration was built to increase the agility of the UAV maneuver and reducing the drag compared to the conventional T-tail. Thus, the need of assessing the accuracy of linearized model occurred to justify its implementation on this unconventional—the V-tail—configuration. This work will perform the investigation of the accuracy of linearized model compared to a more exact numerical solution due to its implementation in unconventional, V-tail configuration. Simulation produced comparative results which conclude that the unconventional V-tail configuration weren’t decreasing the accuracy of the linearized equation compared to complete equation of motion. Thus the linearized model was suitable for the V-tail configuration analysis of its Flight Dynamics.

Propeller design to improve flight dynamics features and performance for coaxial compound helicopters

Ice on aircraft wing changes the aircraft aerodynamic shape, and has negative effects on flight dynamic characteristics, seriously threatening flight safety. Plasma ice shape regulation is a new de-icing method. Plasma actuator produces an apparent thermal effect, which is designed to dissolve the continuous ice into intermittent ice pieces. How to achieve the optimal regulation ice shape to improve the flight dynamics characteristics under icing conditions is a technical problem restricting the application of this method. A simulation ice shape based on previous ice tunnel experiments and a scale model of swept wing were established. The aerodynamic parameters of no ice, full ice, and two regulation ice schemes were obtained based on wind tunnel. Six degrees of freedom flight dynamics model was established, and flight simulation had been carried out. As the analysis of trim characteristics, dynamic stability, and maneuverability, flight dynamics characteristics were better improved when the ratio of ice width to the mean aerodynamic chord was 0.15. The evaluation method of plasma ice shape regulation schemes was proposed. The proposed method, which can compare and optimize the arrangement of plasma actuators, realized the optimal regulation ice shape on the premise of balancing flight safety and energy consumption.

The lightweight structures and unconventional configurations being considered for the next generation of aircraft mean that any effort to predict or control the flight dynamics is impacted by the structural dynamics. One of the most severe forms of coupling between aeroelasticity and flight dynamics is an instability called body freedom flutter. The existing tools often assume a relatively weak effect of structural dynamics on the flight dynamics, and are therefore incapable of modeling strong interactions like body freedom flutter. A method of combining different sources of data traditionally used for aeroelasticity and flight dynamics is described by reconciling many of the differences between these models. By building upon past modeling efforts, a level of familiarity in the approach is achieved. Generally the differences from the traditional approaches are subtle but significant. The traditional frequency domain flutter model in a modal coordinate system is converted to a form consistent with a time domain flight dynamics model. The time domain rational function approximation about a non-inertial coordinate system and the unique constraints for the conversion between the inertial and non-inertial coordinate systems are discussed. A consistent transformation of the states of aeroelastic models to flight dynamics models is derived, which enables the integration of data from higher fidelity computational fluid dynamics models or wind-tunnel testing. The present method of integrating multidisciplinary data was used to create models that compare well with X-56A flight-test data, including conditions past the flutter speed.

The flight dynamics of morphing aircraft must be analyzed to ensure shapechanging trajectories have the desired characteristics. The tools for describing flight dynamics of fixed-geometry aircraft are not valid for time-varying systems such as morphing aircraft. This paper introduces a method to relate the flight dynamics of morphing aircraft by interpreting a time-varying eigenvector in terms of flight modes. The time-varying eigenvector is actually defined through a decomposition of the state-transition matrix and thus describes an entire response through a morphing trajectory. A variable-sweep aircraft is analyzed to demonstrate the information that is obtained through this method and how the flight dynamics are altered by the time-varying morphing. I. Introduction Morphing provides an ability to alter the shape of an aircraft and thus the flight characteristics. Such morphing is particularly attractive for unpiloted air vehicles due to the relative flexibility in their wings. Also, biologically-inspired concepts for morphing are abundant. The flight dynamics of morphing aircraft have received somewhat less attention than the flight performance. In a sense, studies into flight performance consider the characteristics of steadystate morphing whereas studies into flight dynamics consider the characteristics of time-varying morphing. These flight dynamics are critical to understanding the relative coupling between states when maneuvering. The role of eigenvalues and eigenvectors in describing the flight dynamics is well understood for linear time-invariant aircraft. These parameters are directly computed from the state matrix and directly relate properties of the response. Essentially, every real-valued eigenvalue and associated eigenvector describes a divergence or convergence while every complex-conjugate pair of eigenvalues and associated eigenvectors describe an oscillatory mode. This relationship of the flight dynamics with the eigenvalues and eigenvectors results from the exponential solution that is known for the linear time-invariant dynamics. A notion of eigenvalues and eigenvectors is also defined for linear time-varying systems; however, they actually are computed by decomposing the state-transition matrix rather than the state matrix. This requirement notes that the linear time-varying dynamics do not have a closed-form solution so any decomposition of the response must retain the temporal information within the

On 11 September 2001, 19 terrorists overpowered the crews of four commercial passenger aircrafts. The terrorists used the aircraft to commit terrorist attacks by flying the planes into the World Trade Center and in the Pentagon. This form of terrorist attack is described as renegade. Aerial incidents in the following time illustrate that the threat of another renegade is still present. Therefore, the question is still prevailing whether the shooting down of civil aircraft in order to prevent a renegade is lawful. This article is analyzing the legality of shooting down civil aircraft from a public inter- national law perspective. In the first part of the article, it will be discussed that the shooting down of civil aircraft is a violation of the prohibition to use force against civil aircraft in flight which is codified in Article 3bis(a)(1) of the 1944 Convention on International Civil Aviation (so-called Chicago Convention). This norm is also part of customary inter- national law. It will be argued that the limitations of this prohibition, which are codified in Article 3bis(a)(2) of the Chicago Convention are inapplicable as a renegade does not con- stitute an armed attack in the meaning of Article 51 of the Charter of the United Nations – mainly because the terrorist acts cannot be attributed to another State. Limitations of the prohibition to use force against civil aircraft in flight might be also found in other provisions of the Chicago Convention. In the second part of this article, it will be shown that Article 4 of the Chicago Convention prohibits a misuse of civil aviation. Even if the terrorist acts can be attributed to another State and therefore might be consid- ered to be a misuse of civil aviation, Article 4 of the Chicago Convention does not allow the shooting down of civil aircraft. In the third part of this article, limitations of the prohibition to use force against civil aircraft in flight arising out of Article 89 of the Chicago Convention will be examined. In accordance with the findings in the first part, it will be presented that there is regularly no armed conflict, thus Article 89 (1) of the Chicago Convention is inapplicable. Even if there would be an armed conflict, international humanitarian law would still prohibit in most cases the shooting down of civil aircraft as they are regularly civil objects, thus not lawful targets of any armed activity. In addition, the requirements of a necessity are normally not fulfilled which results in an inapplicability of Article 89 (2) of the Chicago Convention. As the shooting down of civil aircraft regularly constitutes a violation of the prohibition to use force against civil aircraft in flight, it is questionable whether this violation is justifiable. The fourth part of this article presents six possible justifications which all arise out of customary international law. These justifications are self defense, countermeasure, force majeur, distress, necessity, and State practice. However, it will be shown that the criteria for each justification are not fulfilled. Therefore, the shooting down of civil aircraft in order to prevent a renegade constitutes regularly an unjustifiable violation of the prohi- bition to use force against civil aircraft in flight.

Flight dynamics modeling and analysis for a Mars helicopter

The emergence of Airbus transformed the market structure of the LCA industry into a duopoly of similar-sized full-range manufacturers. The financing of Airbus's upfront investment expenditures came in a significant proportion from public funds, which violated, in the US's opinion the SCM Agreement. While the Appellate Body follows this view of things to a large extent, it does so in a measured way: the category ofper seillegal export subsidies is interpreted with a view to the manipulation of normal market conditions; the distortion on competitive conditions matters, not the increase of exports as such. Other aspects of subsidies law clarified are the relationship between effect and subsidy. They are closely related but not identical; rightly, the report operates from the premise that the SCM Agreement's regime focuses on the effect, and not on the subsidy as such, which is a manifestation of a political choice by a sovereign Member state. The Appellate Body affirms that a subsidy has a ‘life’, a shorthand for a beginning and an end: it follows that the effect of a subsidy is not bound to be permanent but is bound to terminate. It is to be regretted that the Appellate Body avoided clarifying to what extent partial privatization, hence sale of assets at market prices to private investors, ‘extinguish’ subsidies.

INPE (National Institute for Space Research)’s Satellite Tracking and Control Center (CRC) has been performing TT&C control of satellites since 1993. Currently at CRC, flight dynamics activities and satellite operations planning are performed in an integrated manner. Most payload operations are either performed in real-time, over the visibility range of data receiving ground stations, or else triggered by user requests which conditions are typically predicted by flight dynamics models. Health maintenance required by satellites under CRC control comprises mostly attitude, orbit, power and data handling subsystems telemetry monitoring. Most housekeeping telecommands are planned on a case-by-case basis, upon telemetry trend analysis. The remaining few are mainly onboard ephemeris data updating. Because of these synergies, at CRC the flight dynamics team is also responsible for generating ground and flight operations plans. Software tools currently at use by the CRC flight dynamics team have been proven reliable over decades of continuous running, but are not easy to use or maintain. As technology rapidly changes, new machines must be acquired and, most often, they are not fully compatible with older platforms. Porting program codes from one system to another is costly. Moreover, software operation tasks such as filtering out erroneous measurement data and precisely adjusting input parameters have become a fine art that only the most expert flight dynamics operators can handle. Training a new operator for this purpose alone would be extremely time-consuming. By taking advantage of INPE’s alternative role as an academic institute, CRC has directed some scholarly researches towards satellite control automation. Most resulting works concerned about intelligent planning of flight operations, by managing time and resources among satellite passes over ground stations. Their byproducts would not afford actual spacecraft control operations without major modifications. Due to their academic focusing, they would not reflect the complexity of real systems. In order to fill this gap, a validation tool for such intelligent operations planners is proposed. This project aims at building a software tool that would verify the outputs of academic operations planners by feeding them into a satellite simulator. The tasks contained in the operations plan would be executed on the satellite simulator, which effects would be reflected on its internal statuses. These internal statuses would be compared to design limits set by the satellite developers, in order to either accept or reject the operations plan. It is expected that the resulting validated plan would be workable with actual missions. The final decision on its usability would be decided by the flight dynamics team, upon comparison of results with those from the legacy system currently in use.

An integrated modelling approach, suitable for flight loads analysis and flight dynamics investigations is presented. These scenarios need fast loop capable simulation models. Typically, they are evaluated many times in the entire flight envelope, for varying mass cases and for changing parameters during the aircraft design process. Further, the simulations models have very different requirements depending on their application. Flight dynamics models must account for large nonlinear rigid Body motions and aerodynamic nonlinearities, manoeuvre loads analysis, additionally need to take flexible deformations into account, and hence require distributed aerodynamics, and for gust loads Analysis unsteady aerodynamic effects are essential. The proposed integrated model is based on nonlinear equations of motion using mean axes constraints. Potential flow based aerodynamics with low computational cost are the method of choice for most aeroelastic applications. Presented in this paper are some important aspects regarding the aerodynamic modelling, which are handled differently compared to other implementations. Instead of the Doublet Lattice Method (DLM), the 3D panel method NEWPAN is used, which is able to capture previously neglected flight mechanical effects such as roll-yaw coupling and induced drag. Since, the Aerodynamic Influence Coefficient (AIC) matrices obtained from NEWPAN are nonlinearly dependent on the flight state, a Reduced Order Model (ROM) based on Proper Orthogonal Decomposition (POD) is setup and subsequently interpolated. The proposed ROM interpolates directly AIC matrices (AIC-ROM) instead of pressures. The advantages of a ROM based on AIC matrices over similar approaches based on pressure coefficients are explained. Similarly to the Doublet Lattice Method, the 3D panel code is able to compute frequency based unsteady Aerodynamic Influence Coefficient (AIC) matrices, which are approximated by a Rational Functions (RFA) to make them amenable for time domain simulations. Unlike other approaches, the physical RFA allows a clear separation of quasi-steady, added mass and unsteady lag terms. Several application examples are presented and their particular needs are addressed. The proposed modelling approach supports flight dynamics by including previously neglected aerodynamic roll yaw coupling and nonlinear angle of attack dependencies, manoeuvre loads, including aeroelastic deformations, and gust loads analyses, including unsteady aerodynamics in one integrated model and is suitable for for investigations regarding active gust and manoeuvre load alleviation, primary flight control design, or performance improvement through lift redistribution.

The gas turbine engine is the powerhouse of most large modern military and civil aircraft. These engines operate at temperatures above the melting point of the materials that the combustor and turbine components are made from. Film cooling is used extensively to cool the hot surfaces and extend the life of the gas turbine’s hot end components. In some modern and future engines, the average turbine inlet temperature is increased to about 2,400K and the length of the combustor is reduced. The turbine inlet temperature is increased to improve the thermal efficiency while the combustor is shortened to increase the thrust-to-weight ratio. Both developments are meant to reduce the amount of fuel burnt and the operational cost of the power plant. Increasing the turbine inlet temperature to above 1,850 K CO2 dissociation starts to compete with CO oxidation. Reducing the combustor length reduces the residence time of fuel and increases the likelihood of unburnt hydrocarbons entering the turbine. When carbon monoxide and/or unburnt hydrocarbons enter the turbine, they could react with oxygen in the cooling air and potentially increase the blade metal temperature. An increase of about 30 K can reduce the blade life by half: secondary combustion of reactive species entering the turbine section could therefore lead to serious durability concerns. In a review of the literature, it was found that an estimated 10% of fuel energy is available for combustion in the turbine section and a maximum heat flux augmentation of 18% due to secondary combustion occurs. Secondary combustion in the turbine components is reviewed through a discussion of the analysis of reactive film cooling, developments driving the need to develop an in-depth understanding of reactive film cooling, scaling of reaction kinetics and heat release potential, performance of cooling hole geometries and configurations and mitigation strategies.