Showing papers on "Rocket published in 2016"

Injector-Driven Combustion Instabilities in a Hydrogen/Oxygen Rocket Combustor

Abstract: Self-excited combustion instabilities of the first tangential mode have been found in a research combustor operated with the cryogenic propellant combination of hydrogen/oxygen. In a series of consecutive test campaigns, the influence of operating conditions on these self-excited combustion instabilities was examined. This included a variation of the combustion chamber pressure, the mixture ratio, and the propellant temperatures. It has been shown how these operating parameters influence the resonance frequencies of the combustion chamber. The analysis of the influence of operating conditions on the oscillation amplitude of the first tangential mode indicated that the instability occurred when the frequency of the first tangential mode of the combustion chamber was shifted into the frequency of the second longitudinal mode of the liquid oxygen injector. With a variation of the injector length, and therefore its longitudinal resonance frequencies, this hypothesis has been tested. Based on the experimental ...

A compact and robust diode laser system for atom interferometry on a sounding rocket

Abstract: We present a diode laser system optimized for laser cooling and atom interferometry with ultra-cold rubidium atoms aboard sounding rockets as an important milestone toward space-borne quantum sensors. Design, assembly and qualification of the system, combing micro-integrated distributed feedback (DFB) diode laser modules and free space optical bench technology, is presented in the context of the MAIUS (Matter-wave Interferometry in Microgravity) mission. This laser system, with a volume of 21 l and total mass of 27 kg, passed all qualification tests for operation on sounding rockets and is currently used in the integrated MAIUS flight system producing Bose–Einstein condensates and performing atom interferometry based on Bragg diffraction. The MAIUS payload is being prepared for launch in fall 2016. We further report on a reference laser system, comprising a rubidium stabilized DFB laser, which was operated successfully on the TEXUS 51 mission in April 2015. The system demonstrated a high level of technological maturity by remaining frequency stabilized throughout the mission including the rocket’s boost phase.

Hybrid Rockets with Axial Injector: Port Diameter Effect on Fuel Regression Rate

Abstract: A numerical rebuilding of a series of test data obtained with static firings of a lab-scale hybrid rocket is carried out with a Reynolds-averaged Navier–Stokes solver including detailed gas–surface interaction modeling based on surface mass and energy balances. Two experimental campaigns are considered in which gaseous oxygen is fed into axisymmetric hydroxyl-terminated polybutadiene grains through an axial conical subsonic nozzle. A validation rebuilding of all of the firing tests has been performed first to highlight numerical prediction capabilities and modeling limits. Despite the several geometrical simplifications, which allows using a reduced number of cells in the computational domain and thus performing parametric analyses efficiently, the present computational fluid dynamics approach is able to capture the main features of the motor internal ballistics, fairly reproducing the average chamber pressure values and the fuel regression rate trends with oxidizer mass flux and port diameter. Computed f...

Fundamentals of Aircraft and Rocket Propulsion

Abstract: This book provides a comprehensive basics-to-advanced course in an aero-thermal science vital to the design of engines for either type of craft. The text classifies engines powering aircraft and single/multi-stage rockets, and derives performance parameters for both from basic aerodynamics and thermodynamics laws. Each type of engine is analyzed for optimum performance goals, and mission-appropriate engines selection is explained. Fundamentals of Aircraft and Rocket Propulsion provides information about and analyses of: thermodynamic cycles of shaft engines (piston, turboprop, turboshaft and propfan); jet engines (pulsejet, pulse detonation engine, ramjet, scramjet, turbojet and turbofan); chemical and non-chemical rocket engines; conceptual design of modular rocket engines (combustor, nozzle and turbopumps); and conceptual design of different modules of aero-engines in their design and off-design state. Aimed at graduate and final-year undergraduate students, this textbook provides a thorough grounding in the history and classification of both aircraft and rocket engines, important design features of all the engines detailed, and particular consideration of special aircraft such as unmanned aerial and short/vertical takeoff and landing aircraft. End-of-chapter exercises make this a valuable student resource, and the provision of a downloadable solutions manual will be of further benefit for course instructors

A three-layer magnetic shielding for the MAIUS-1 mission on a sounding rocket

Abstract: Bose-Einstein-Condensates (BECs) can be used as a very sensitive tool for experiments on fundamental questions in physics like testing the equivalence principle using matter wave interferometry. Since the sensitivity of these experiments in ground-based environments is limited by the available free fall time, the QUANTUS project started to perform BEC interferometry experiments in micro-gravity. After successful campaigns in the drop tower, the next step is a space-borne experiment. The MAIUS-mission will be an atom-optical experiment that will show the feasibility of experiments with ultra-cold quantum gases in microgravity in a sounding rocket. The experiment will create a BEC of 105 87Rb-atoms in less than 5 s and will demonstrate application of basic atom interferometer techniques over a flight time of 6 min. The hardware is specifically designed to match the requirements of a sounding rocket mission. Special attention is thereby spent on the appropriate magnetic shielding from varying magnetic fields during the rocket flight, since the experiment procedures are very sensitive to external magnetic fields. A three-layer magnetic shielding provides a high shielding effectiveness factor of at least 1000 for an undisturbed operation of the experiment. The design of this magnetic shielding, the magnetic properties, simulations, and tests of its suitability for a sounding rocket flight are presented in this article.

Experimental and Numerical Investigation of a Solid-Fuel Rocket Scramjet Combustor

Abstract: A new scramjet configuration using solid fuel as propellant is proposed, namely, the solid-fuel rocket scramjet. Experimental and numerical investigation of the solid-fuel rocket scramjet combustor was conducted to evaluate its performance. The experiment simulated a flight environment of Mach 4 at a 17 km altitude. Magnesium-based solid fuel was used as propellant in this study. The results show that secondary combustion occurs in supersonic combustor. The combustion efficiency of the propellant is about 65%, and the total pressure recovery is about 0.5. These imply that the scramjet configuration using solid fuel as a propellant is feasible. The results of the numerical analysis show that a normal shock wave is generated in the core flow of the combustor, and this is due to all of the fuel-rich gas being injected in the same cross section.

Reusable Rocket Engine: Firing Tests and Lifetime Analysis of Combustion Chamber

Abstract: The main engine of the reusable sounding rocket has been developed at the Kakuda Space Center of the Japan Aerospace Exploration Agency. A total of 54 engine firing experiments were conducted between June 2014 and February 2015. Throughout these experiments, the advanced capabilities and functions of this engine, such as wide-range throttling, accurate controllability, and health monitoring, were proved to be feasible. This rocket was also required to be reusable for over 100 flights. To confirm the long-life durability of this engine for over 100 flights in a limited experiment period, sequential multiple-firing tests were planned and conducted. Numerical analysis to evaluate damage of the main chamber due to firings was also conducted. The results of the durability tests and numerical analysis showed reuse of this engine for over 100 flights to be feasible.

Minimum Time Control of the Rocket Attitude Reorientation Associated with Orbit Dynamics

Abstract: In this paper, we investigate the minimal time problem for the guidance of a rocket, whose motion is described by its attitude kinematics and dynamics but also by its orbit dynamics. Our approach is based on a refined geometric study of the extremals coming from the application of the Pontryagin maximum principle. Our analysis reveals the existence of singular arcs of higher order in the optimal synthesis, causing the occurrence of a chattering phenomenon, i.e., of an infinite number of switchings when trying to connect bang arcs with a singular arc. We establish a general result for bi-input control-affine systems, providing sufficient conditions under which the chattering phenomenon occurs. We show how this result can be applied to the problem of the guidance of the rocket. Based on this preliminary theoretical analysis, we implement efficient direct and indirect numerical methods, combined with numerical continuation, in order to compute numerically the optimal solutions of the problem.

Numerical Investigation of Flow and Combustion in a Single Element GCH4/GOx Rocket Combustor

Abstract: The flow and combustion in a GCH4/GOX single-element rocket combustor is analysed by several groups using different numerical models and tools. The tools and simulation setups vary with respect to modeling fidelity and computational expense. A short overview of the tools and the individual simulation setups is given. The focus of the paper is the comparison of the results obtained by the different groups as well as with experimental data. This encompasses the study of specific features of the combustor flow among the different simulations, as well as the validation with typical rocket engine design and performance parameters, such as wall heat flux and combsution pressure, gained from hot firing tests.

The energetic potential of azo- and azoxyfurazan nitro derivatives as components of composite rocket propellants

Abstract: The energetic potential of solid composite rocket propellants on the basis of solid and liquid nitro derivatives of azo- and azoxyfurazans was studied by using thermochemical calculations. Quantitative relationships were established between the energetic parameters of propellants, the selected binder, and the properties of the studied component. It was shown that the studied solid nitrofurazans can serve as the foundation of metal-free compositions of solid composite rocket propellants, giving specific impulse over 263 s, while the use of liquid nitrofurazans as plasticizers of the binder can improve the ballistic effectiveness of rocket propellants containing aluminum hydride by 2–4 s.

Flight Validation of a Rotary-Valved Four-Cylinder Pulse Detonation Rocket

Abstract: A rotary-valved four-cylinder pulse detonation rocket engine system, Todoroki II, was developed, in which two novel techniques, the use of an inflow-driven motor and an inverted oxidizer cylinder, were introduced. The total length of the system was 1910 mm; its total weight when filled with ethylene–nitrous-oxide propellant and helium purge gas was 32.5 kg; and the engine weight was 9.6 kg. In a ground firing test with a duration of 1500 ms, a thrust-to-engine-weight ratio of 2.7 was achieved. Thus, it was demonstrated that a multicylinder pulse detonation rocket engine system can be used as a practical thrust mechanism. Using a launch and recovery system, a flight-simulating test was conducted to evaluate the features and viability of the engine design. The launch and recovery system operated perfectly, and Todoroki II reached a height of about 9.7 m. The operation of the pulse detonation rocket engine under conditions simulating real vertical flight without constraint forces with a duration of about 120...

Daytime midlatitude plasma depletions observed by Swarm: Topside signatures of the rocket exhaust

Abstract: The daytime midlatitude plasma depletions (DMLPDs) observed on 22 May 2014 and 20 May 2015 by the Swarm constellation are not explained by any known natural phenomena. The DMLPDs were detected after rocket launches, and the DMLPD traces converged to the launch station. The event in 2015, for which sufficient total electron content (TEC) data are available, is accompanied with TEC depletion lasting for about 6 h. The persistence generally agrees with the lifetime expected for rocket exhaust depletions (REDs) which is determined by the recombination of the ionospheric oxygen ion with water molecules in the rocket exhaust. These results lead to the conclusion that DMLPDs are REDs in the topside. The RED characteristics identified from the observations on both days are (1) enhancement in electron temperature, (2) reduction in electron pressure, and (3) absence of substructures down to scale sizes of about 8 km (Nyquist’s scale size).

Some Aspects on Safety and Environmental Impact of the German Green Gel Propulsion Technology

Abstract: This article gives a short introduction into the technology of gelled propellant rocket motors and gas generators. A brief introduction outlines the specific Features of the German green gelled propellant rocket Motor technology and the family of monopropellants that cover a variety of requirements with respect to smoke, combustion temperature, and combustion pressure ranges. The discussion of the hazard potential comprises military insensitive munitions (IM) and civil classifications as well as a comparative assessment of the environmental and health Impacts from manufacturing over use and finally to disposal. Summarizing the properties over all categories shows that gelled propellants provide a unique combination of good insensitivity and low environmental and health hazard potential compared to other liquid propellants, fuels, and oxidizers or solid rocket propellants.

Validation of Conjugate Heat Transfer Model for Rocket Cooling with Supercritical Methane

Abstract: A numerical solver able to describe a rocket engine cooling channel fed with supercritical methane is validated against experimental data coming from a test article conceived and tested by the Italian Aerospace Research Center. The multidimensional conjugate heat transfer model numerically solves the Reynolds-averaged Navier–Stokes equations for the coolant flow and the Fourier’s law of conduction for the heat transfer within the wall. In this study, an experimental test case is reproduced in detail in order to evaluate the influence of partially unknown parameters, such as surface roughness and wall thermal conductivity, and of operative parameter uncertainty, such as the coolant mass flow rate and input heat transfer rate. The comparison made with respect to the wall temperature and coolant pressure drop of the whole set of experimental data provides complementary information that allows better understanding of experiments and infers possible deviations from the expected behavior.

Vertical returning trajectory design method for carrier rocket

Abstract: The invention provides a vertical returning trajectory design method for a sub-stage of a carrier rocket. The method comprises: according to flight characteristics of a sub-stage of a carrier rocket, determining a vertical returning launching point or each sub-flight-phase of a vertical returning preset target position, using a three-freedom-degree kinetic model of a sub-stage vertical returning process on an earth spherical model, to generate a vertical returning ballistic trajectory, wherein each sub-flight-phase from sub-stage separation to sub-stage landing includes a posture alignment stage, a deceleration turning stage, a gliding stage, a power deceleration stage, a pneumatic deceleration stage, a vertical dropping stage, or a gliding posture alignment stage, a power deceleration stage, a deceleration stage, and a vertical dropping stage, the kinetic model being determined by flight speed, a trajectory inclination angle, a trajectory rotation angle, speed and position components in a launching coordinate system, an attack angle, a sideslip angle, and a variable thrust factor. The invention also provides a vertical returning trajectory design method for a carrier rocket booster stage and a carrier rocket. The vertical returning trajectory design methods are simple in operation and easy in engineering implementation. Track indexes of each obtained sub-flight-phase satisfy constraint requirements of heat flux peak, dynamic pressure, flight overload, and terminal positions.