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ROCKET and ramjet engines have not the universal application that gas turbines command and possibly on this account they have not had, until recent years, the development effort which gave such amazing results in turbine powered aircraft. Nevertheless, they have demonstrated quite dramatically in various parts of the world that they are power plants to be reckoned with. In Great Britain, their value for aircraft was appreciated somewhat belatedly and events have since decreed that the promise they showed should be smothered before it could become a vital fact. On the other hand their importance for missiles was realized at the conclusion of the 1939–45 war, but again they were not encouraged on anything like the scale that present events show would have been justified. Because of this lack of encouragement, British rockets and ramjets, instead of leading the world, as do gas turbines, are struggling hard to provide a modest rate of progress.

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.

The purpose of this paper is to attempt an aerospaceplane design with the objective of Low-Earth-Orbit-and-Return-to-Earth (LEOARTE) under the constraints of safety, low cost, reliability, low maintenance, aircraft-like operation and environmental compatibility. Along the same lines, a “sister” point-to-point flight on Earth Suborbital Aerospaceplane is proposed.

The LEOARTE aerospaceplane is based on a simple design, proven low risk technology, a small payload, an aerodynamic solution to re-entry heating, the high-speed phase of the outgoing flight taking place outside the atmosphere, a propulsion system comprising turbojet and rocket engines, an Air Collection and Enrichment System (ACES) and an appropriate mission profile.

It was found that a LEOARTE aerospaceplane design subject to the specified constraints with a cost as low as 950 United States Dollars (US$) per kilogram into Low Earth Orbit (LEO) might be feasible. As indicated by a case study, a LEOARTE aerospaceplane could lead, among other activities in space, to economically viable Space-Based Solar Power (SBSP). Its “sister” Suborbital aerospaceplane design could provide high-speed, point-to-point flights on the Earth.

The proposed LEOARTE aerospaceplane design renders space exploitation affordable and is much safer than ever before.

This paper provides an alternative approach to aerospaceplane design as a result of a new aerodynamically oriented Thermal Protection System (TPS) and a, perhaps, improved ACES. This approach might initiate widespread exploitation of space and offer a solution to the high-speed “air” transportation issue.

THE effect of the propulsive efficiency is analysed for all speed regions and methods for obtaining optimum propulsive efficiency for any speed and environmental conditions are investigated for different type vehicles. Also, the importance of the propulsive efficiency is compared to other factors such as weight of power plant, specific fuel consumption, specific power ratio, specific thrust ratio, etc. Finally, on the basis of considering all power plant factors, it is shown how to achieve optimum propulsion for any vehicle at required operating conditions.

The purpose of this study is to determine the impact of two parallel boosters fixed to the ILR 33 AMBER 2 K core rocket stage on its aerodynamic characteristics in the subsonic and transonic regimes and for M = 2.3.

Wind tunnel tests of the rocket model were carried out in a trisonic wind tunnel using a six-component internal balance. Three rocket model configurations were investigated.

The results of the presented studies showed that the presence of boosters causes a significant increase in the total rocket drag, which depends on both the Mach number and the rocket flight phase. Experimental tests of the rocket model allowed to determine the difference in drag coefficient between active and passive flight versus Mach number. It was found that, in the case of a deviation from the rocket’s flight direction, the aerodynamic coefficients strongly depend on the location of the boosters in relation to the direction of the deviation.

Studies of the rocket model aerodynamic characteristics allow the assessment of the influence of parallel boosters on rocket performance, which is important when the decision of a rocket staging type is taken.

The presented wind tunnel test results of the rocket model equipped with the two parallel boosters are an original contribution to the rocket research results presented in the literature.

THE use of metals at temperatures in excess of 1,200 deg. F. and up to temperatures in the vicinity of their melting points is a challenging and fascinating portion of the fight to pass the heat barrier in the design and performance of aircraft and their power plants. The materials available for service in this temperature range are restricted. The considerations of designing structural components involve many more problems than the old criteria of strength to weight ratio and fabrication costs. Such properties as thermal expansion, heat conductivity, surface emissivity and scaling resistance are as important in determining which metal should be used for a given application as are the various measurements of strength heretofore the primary considerations in material selection.

THE use of liquid oxygen as an oxidizer for various fuels in liquid rocket propellent systems is not new. Professor Goddard used liquid oxygen in his rocket experiments and the well known German V‐2 rockets used this material as an oxidizer. However, its effect on non‐metallic materials ordinarily used in rocket systems was not investigated until recent years. This investigation was prompted by phenomena which had been experienced by rocket engine and rocket aircraft manufacturers and by suppliers of the material. It was observed that when some organic materials came in intimate contact with liquid oxygen they became prone to detonation when subjected to certain impact energies. This was undoubtedly due to the formation of unstable organo‐peroxide compounds which when impacted released high levels of energy resulting in an explosion. Specifically, when liquid oxygen was accidently spilled on asphalt and inadvertently stepped on, the asphalt would often explode. Also, leather gaskets immersed in liquid oxygen and subjected to surge impact detonated with disastrous effects.

A new concept solution for improving blast survivability of the light tactical military vehicles is proposed and critically assessed using computational engineering methods and tools.

The solution is inspired by the principle of operation of the rocket-engine nozzles, in general and the so called “pulse detonation” rocket engines, in particular, and is an extension of the recently introduced so-called “blast chimney” concept (essentially a vertical channel connecting the bottom and the roof and passing through the cabin of a light tactical vehicle). Relative to the blast-chimney concept, the new solution offers benefits since it does not compromise the cabin space or the ability of the vehicle occupants to scout the environment and, is not expected to, degrade the vehicle’s structural durability/reliability. The proposed concept utilizes side vent channels attached to the V-shaped vehicle underbody whose geometry is optimized with respect to the attainment of the maximum downward thrust on the vehicle. In the course of the channel design optimization, analytical and computational analyses of supersonic flow (analogous to the one often used in the case of the pulse detonation engine) are employed.

The preliminary results obtained reveal the beneficial effects of the side channels in reducing the blast momentum, although the extent of these effects is quite small (2-4 per cent).

To the authors’ knowledge, the present work is the first exploration of the side-vent-channels concept for mitigating the effect of buried-mine explosion on a light tactical vehicle.