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The purpose of this paper is to present a novel approach in aeroservoelastic analysis and robust control of a wing section with two control surfaces in leading‐edge and trailing‐edge. The method demonstrates how the number of model uncertainties can affect the flutter margin.

The proposed method effectively incorporates the structural model of a wing section with two degrees of freedom of pitch and plunge with two control surfaces on trailing and leading edges. A quasi‐steady aerodynamics assumption is made for the aerodynamic modeling. Basically, perturbations are considered for the dynamic pressure models and uncertainty parameters are associated with structural stiffness and structural damping and are accounted for in the model by a Linear Fractional Transformation (LFT) model. The control commands are applied to a first and second order electro‐mechanical actuator.

Dynamic performance of aeroelastic/aeroservoelastic system including time responses, system modal specifications, critical flutter speeds, and stability margins are extracted and compared with each other. Simulation results are validated through experiments and are compared to other existing methods available to the authors. Results of simulations with four structural uncertainties and first order controllers have a good agreement with experimental test results. Furthermore, it is shown that by using a high gain second order controller, the aeroservoelastic (ASE) system does not have any coupling nature in frequency response.

In this study, modeling, simulation, and robust control of a wing section have been investigated utilizing the μ‐Analysis method and the wing flutter phenomenon is predicted in the presence of multiple uncertainties. The proposed approach is an advanced method compared to conventional flutter analysis methods (such as V‐g or p‐k) for calculating stability margin of aeroelastic/aeroservoelastic systems.

IN Parts I, II and III of this series we have discussed the physical nature of divergence, control reversal and various forms of flutter, and have seen how these phenomena can be predicted by theory. The flutter problem is so complicated, however, that the aircraft designer needs the assistance of certain guiding principles; otherwise he may find when the aircraft is ready to fly that the flutter calculations which are just completed show that drastic modifications to the aircraft are necessary. These principles form the basis of this concluding part of the series and have two main objects: first to avoid large changes in design on flutter grounds and secondly to obtain a high efficiency from the flutter calculations.

BEFORE a detailed consideration of internal stresses may be made, it is necessary to define external loadings which are possibly critical. This involves the consideration of manoeuvres throughout the altitude range of the aeroplane, to a severity fixed by aerodynamic or specification values of speed and normal acceleration.

THE Annual Meeting proper was preceded by the Honors Dinner at which several awards were given for distinguished achievements in aeronautics. The Daniel Guggenheim Medal itself was presented to Glenn L. Martin, an early aviation pioneer, whose Baltimore plant is turning out patrol bombers, the B‐26, an advanced medium bomber, and the Baltimore's medium bomber specially designed for British needs. Mr. Martin is the last survivor of the American aeroplane pioneers to head an aircraft manufacturing company bearing his own name. William J. Knudsen was among the distinguished guests. The guest of honour was Griffith Brewer, President of the Royal Aeronautical Society of Great Britain, who called for speedy help to Britain in its hour of need. If the enthusiastic greeting given to Mr. Brewer is any criterion, then it may be taken for granted that the aviation fraternity of the United States is heart and soul with the British stand.

The ultimate aim of the stress calculation of aircraft structures is to reduce the frequency of defects in service to an acceptably low level. If this is to be done without undue structure weight the design loads, the factors of safety and the allowable stresses must be chosen with great care. In principle, there must be some relation between the probability of failure and the design strength. On a new design this function is always unknown and the designer must rely on experience of previous aircraft to guide his judgment. However, the required service life and the expected conditions of service, including temperature effects, vibration, etc., must be foreseen and taken into consideration.

A study of the world altitude records for heavier‐than‐air aircraft homologated by the Fédération Aéronautique Internationale is interesting, and yields the information that from 1909 to 1914 the mean rate of increase of ceiling was almost 4,500 ft. per year, while from 1920 to 1936 the mean rate dropped to 1,000 ft. per year. Slumps, depressions and other financial phenomena must obviously have a repercussive influence in this particular field of aircraft development, and a pertinent example of their effect is provided by the unfortunate hiatus which occurs from 1914 to 1920, and obscures the height at which the change in rate takes place. It has been shown by Mr. McKinnon Wood, in an unpublished paper on the design of an aeroplane to reach a great height, written in November, 1931, that the power which the engine must develop at ceiling increases as the ceiling is raised, and that, in the case of a supercharged engine, the difference between the supercharged height and the ceiling of the aircraft decreases as the ceiling increases. Thus, to maintain the initial rate of increase of ceiling an accelerating rate of progress in engine design was necessary, but unlikely in view of the increasing number and difficult nature of the problems to be encountered and solved.

The purpose of this research is to explore innovative aircraft concepts that use flexible wings and distributed propulsion to significantly reduce fuel burn of future transport aircraft by exploiting multidisciplinary interactions.

Multidisciplinary analysis and trajectory optimization are used to evaluate the mission performance benefits of flexible wing distributed propulsion aircraft concepts.

The flexible wing distributed propulsion aircraft concept was shown to achieve a 4 per cent improvement in L/D over a mission profile consisting of a minimum fuel climb, minimum fuel cruise and continuous descent.

The technologies being investigated may lead to mission adaptive aircraft that can minimize drag, and thus fuel burn, throughout the flight envelope.

The aircraft concepts being explored seek to create synergistic interactions between disciplines for reducing fuel burn while capitalizing on the potential benefits of lightweight, flexible wing structures and distributed propulsion.

The paper aims to present a design and numerical verification procedure of a composite casing of a microstrip antenna for an aerospace satellite.

The casing for the microstrip antenna was designed in a form of a laminate shell with variable number of layers of reinforcing fabric. The material properties, both static and dynamic, were determined experimentally and then exported to an environment of numerical analyses. The numerical modal analysis allows optimizing the geometry and lay-up of the casing in such a way that a number of modal shapes occurring in the operational frequency band was significantly reduced, several modal shapes with high displacement in flanges of the casing were eliminated and the values of natural frequencies were increased. A final model of the composite casing was subjected to two types of analyses which simulate typical operation conditions during spacecraft mission. These analyses contained thermomechanical quasi-static analyses with 12 loadcases and thermomechanical shock analyses with 9 loadcases, which simulate various mechanical and temperature conditions.

Results of the performed analyses were compared with safety margins determined by following requirements to spacecraft vehicles. The obtained results confirm the design feasibility, which allow considering the proposed design during manufacturing of a prototype in further studies.

Moreover, the presented results can be considered as a design methodology guideline, which can be helpful for engineers working in the aerospace industry.

The originality of the paper lies in the proposed design and verification procedure of composite elements subjected to operational loading during a spacecraft mission.

Since Hayek’s pioneering work in the 1930s, the Austrian business cycle theory (ABCT) has been presented as a disequilibrium theory populated by less-than-perfectly rational agents. In contrast, we maintain that (1) the Austrian business cycle theory is consistent with rational expectations and (2) the post-boom adjustment process can be understood in an equilibrium framework. Hence, we offer a new interpretation of the existing theory. In doing so, we also address concerns raised with Garrison’s (2001) diagrammatic approach, wherein the economy moves beyond the production possibilities frontier. Our interpretation might accurately be described as a monetary disequilibrium approach grounded in an implicit general equilibrium framework with positive costs of reallocation.

Today’s modern liquid propellant rocket engines have a very complicated structure. They cannot be arbitrarily downsized, ensuring efficient propellants’ mixing and combustion. Moreover, the thermodynamic cycle’s efficiency is relatively low. Utilizing detonation instead of deflagration could lead to a significant reduction of engine chamber dimensions and mass. Nowadays, laboratory research is conducted in the field of rotating detonation engine (RDE) testing worldwide. The aim of this paper is to cover the design of a flight demonstrator utilizing rocket RDE technology.

It presents the key project iterations made during the design of the gaseous oxygen and methane-propelled rocket. One of the main goals was to develop a rocket that could be fully recoverable. The recovery module uses a parachute assembly. The paper describes the rocket’s main subsystems. Moreover, vehicle visualizations are presented. Simple performance estimations are also shown.

This paper shows that the development of a small, open-structure, rocket RDE-powered vehicle is feasible.

Flight propulsion system experimentation is on-going. However, first tests were conducted with lower propellant feeding pressures than required for the first launch.

Importantly, the vehicle can be a test platform for a variety of technologies. The rocket’s possible further development, including educational use, is proposed.

Up-to-date, no information about any flying vehicles using RDE propulsion systems can be found. If successful in-flight experimentation was conducted, it would be a major milestone in the development of next-generation propulsion systems.