Search results

The purpose of this paper is to present the aerodynamic analysis and external ballistics modeling used in the development of a rocket-target for short range air defence missile systems.

A computational fluid dynamics (CFD) analysis of the airflow around the rocket-target was carried out to estimate the drag, which was needed to develop a mathematical model for external ballistics of the rocket-target. Field-experimental testing was conducted to compare the model results to the data obtained experimentally using various additional measurement techniques such as global positioning system (GPS) coordinates marking of the crash and launch sites, air defence surveillance radar tracking and installing equipment for telemetric data capturing and transmission.

Various ballistic parameters such as the velocity and trajectory of the rocket-target were obtained taking into account the CFD analysis results and internal ballistics data. The field-experimental testing showed a good agreement between the model results and the results obtained by the experimental techniques.

The presented computational models and the experimental techniques could be used in future developments of similar aircraft.

This paper presents a research approach for developing a rocket-target. The results of the research were used as a basis for developing a rocket-target for short range air defence rocket systems. The developed rocket-target was successfully implemented in practice.

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.

THOUGH rockets have aroused a good deal of public interest during the last few years and a great number of very interesting books and articles have been published about the theoretical side of this new science, little is generally known about the experimental progress that has been made, especially in Germany and the U.S.A. In describing this science—the Americans call it “rocketry”—as “new,” it is to be understood that this term applies only to the mathematics of it. The ordinary powder or “sky” rocket is by no means new, but has a long and very involved history, going back to Hassan Alrammah, called “nedshm‐eddin” (The Faithful) in A.D. 1280, who designed the first rocket‐driven torpedo. But though rockets in general (i.e. the powder rocket, which alone existed previous to 1929) have a history of almost a millennium and have even been of historical importance (Sir William Congreve's war‐rockets), the manufacturers of powder rockets knew nothing about their mathematics. When, for example, in 1928 the German Verein für Raumschiffahrt discussed the problem of exhaust velocities and impulses, its president, Johannes Winkler, asked the largest rocket factories about this information and received the answer that they did not know it and had no way of determining it. Winkler was therefore obliged to take the thrust‐diagram of a powder rocket himself (Fig. 1). This diagram revealed that the thrust of a sky‐rocket lasts for only two‐tenths of a second; this result was really amazing and the most amazed were the manufacturers of these rockets.

UNDER the auspices of the Gassiot Com‐mittee of the Royal Society a conference on rocket exploration of the upper atmosphere was held at Queen's College, Oxford, on August 24, 25 and 26. The majority of the papers pre‐sented were concerned with the instrumentations and techniques for physical measurements con‐cerning the atmosphere, but in addition there were some papers which outlined the develop‐ment of the high altitude sounding rocket, first using modified V‐2 rockets and then new vehicles constructed specifically for high altitude research.

The aim of the research is to conduct a study into a configuration of an aircraft system with a focus on aerodynamics. In addition, trim condition and static stability constraints were included. The main application of this system is suborbital space flights. The presented concept of a modular airplane system (MAS) consists of two vehicles: a Rocket Plane and a Carrier. Both are designed in tailless configurations but coupled formed a classic tail aircraft configuration, where the Rocket Plane works as the empennage. The most important challenge is to define the mutual position of those two tailless vehicles under the assumption that each vehicle will be operating alone in different flight conditions while joined in one object create a conventional aircraft. Each vehicle configuration (separated and coupled) must fulfil static stability and trim requirements.

Aircrafts’ aerodynamic characteristics were obtained using the MGAERO software which is a commercial computing fluid dynamics tool created by AMI Aero. This software uses the Euler flow model. Results from this software were used in the static stability and trim condition analysis.

The main outcome of this investigation is a mutual position of the Rocket Plane and the Carrier that fulfils project requirements. Also, the final configuration of both separated vehicles (Rocket Plane and Carrier) and the complete MAS were defined. In addition, it was observed that in the case of classic aircraft configuration which is created by connecting two tailless vehicles increasing horizontal tail arm reduces static stability. This is related to a significantly higher mass ratio of the horizontal tail (the Rocket Plane) with respect to the whole system. Moving backward, the Rocket Plane has a notable effect on a position of a centre of gravity of the whole system static stability. Moreover, the impact of the mutual vehicles’ position (horizontal tail arm) and inclination angle on the coupled vehicle lift to drag ratio was analysed.

In terms of aerodynamic computation, MGAERO software using an inviscid flow model, therefore, both a friction drag and breakdown of vortex are not considered. But the presented research is for the computation stage of the design, and the MGAERO software guarantees satisfactory accuracy with respect to the relatively low time of computations. The second limitation is that the presented results are for the conceptual stage of the design and dynamic stability constraints were not taken into account.

The ultimate goal of the coupled aircraft project is to conduct flying tests and the presented result is one of the milestones to achieve this goal.

A design process for a conventional aircraft configuration is well known however, there are not many examples of vehicles that consist of two coupled aircrafts where both vehicles have similar mass. The unique part of this paper includes results of the investigation of the mutual position of the vehicles that can fly alone, as well as in coupled form. The impact of the position of the centre of gravity on trim conditions and static stability of the coupled configuration was investigated.

According to the study of the space flight market, there is a demand for space suborbital flights including commercial tourist flights. However, one of the challenges is to design a mission and a vehicle that could offer flights with relatively low G-loads. The project of the rocket-plane in a strake-wing configuration was undertaken to check if such a design could meet the FAA recommendation for this kind of flight. The project concept assumes that the rocket plane is released from a slowly flying carrier plane, then climbs above 100 kilometers above sea level and returns in a glide flight using a vortex lift generated by the strake-wing configuration. Such a mission has to include a flight transition during the release and return phases which might not be comfortable for passengers. Verification if FAA recommendation is fulfilled during these transition maneuvers was the purpose of this study.

The project was focused on the numerical investigation of a possibility to perform transition maneuvers mentioned above in a passenger-friendly way. The numerical simulations of a full-scale rocket-plane were performed using the simulation and dynamic stability analyzer (SDSA) software package. The influence of an elevator deflection change on flight parameters was investigated in two cases: a transition from the steep descent at high angles of attack to the level glide just after rocket-plane release from the carrier and an analogous transition after re-entry to the atmosphere. In particular, G-loads and G-rates were analyzed.

As a result, it was found that the values of these parameters satisfied the specific requirements during the separation and transition from a steep descent to gliding. They would be acceptable for an average passenger.

To verify the modeling approach, a flight test campaign was performed. During the experiment, a rocket-plane scaled model was released from the RC model helicopter. The rocket-plane model was geometrically similar only. Froude scales were not applied because they would cause excessive technical complications. Therefore, a separate simulation of the experiment with the application of the scaled model was performed in the SDSA software package. Results of this simulation appeared to be comparable to flight test results so it can be concluded that results for the full-scale rocket-plane simulation are also realistic.

It was proven that the rocket-plane in a strake-wing configuration could meet the FAA recommendation concerning G-loads and G rates during suborbital flight. Moreover, it was proven that the SDSA software package could be applied successfully to simulate flight characteristics of airplanes flying at angles of attack not only lower than stall angles but also greater than stall angles.

The application of rocket-planes in a strake-wing configuration could make suborbital tourist flights more popular, thus facilitating the development of manned space flights and contributing to their cost reduction. That is why it was so important to prove that they could meet the FAA recommendation for this kind of service.

The original design of the rocket plane was analyzed. It is equipped with an optimized strake wing and is controlled with oblique, all moving, wingtip plates. Its post-stall flight characteristics were simulated with the application of the SDSA software package which was previously validated only for angles of attack smaller than stall angle. Therefore, experimental validation was necessary. However, because of excessive technical problems caused by the application of Froude scales it was not possible to perform a conventional test with a dynamically scaled model. Therefore, the geometrically scaled model was built and flight tested. Then a separate simulation of the experiment with the application of this model was performed. Results of this separate simulation were compared with the results of the flight test. This comparison allowed to draw the conclusion on the applicability of the SDSA software for post-stall analyzes and, indirectly, on the applicability of the proposed rocket-plane for tourist suborbital flights. This approach to the experimental verification of numerical simulations is quite unique. Finally, a quite original method of the model launching during flight test experiment was applied.

ALTHOUGH a new epoch in high‐speed flight opened with the introduction of the jet propulsion engine, new problems were presented by the thrust characteristics and high fuel‐consumption of the gas turbine.

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 rocket‐propelled aeroplane demands the successful combination of two conflicting principles. In fact the problem has been compared with the construction of a ship equally suited for travelling on rivers and navigating the high seas. The principle of the rocket, as that of the aeroplane, is to move by reaction against an accelerated mass. There is a great difference, however, in the manner of application of this principle, as the following two examples will show. A powder rocket of a commercial type has an overall, weight of 140 grammes (453 grammes = 1 lb.); the weight of the shell containing the driving charge is 40 grammes, while the charge weighs approximately 15 grammes. The effective burning time of such a rocket was found to be 0.15 sec.; the thrust during this period being 4.8 legs, or approximately 11 lb. The exhaust speed of the gases (“c”) was calculated to be 453 metres per second; the acceleration 314 metres per second (in other words more than 31 gravities): the altitude 110 metres without considering air resistance. Actually the altitude reached was about 80 metres, or 250 ft.

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.