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Purpose 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. Design/methodology/approach 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. Findings This paper shows that the development of a small, open-structure, rocket RDE-powered vehicle is feasible. Research limitations/implications Flight propulsion system experimentation is on-going. However, first tests were conducted with lower propellant feeding pressures than required for the first launch. Practical implications Importantly, the vehicle can be a test platform for a variety of technologies. The rocket’s possible further development, including educational use, is proposed. Originality/value 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.

Accepted: 2015-09-15, 資料番号: SA1150300000

In this paper, an in-flight prediction method of thrust profiles for solid rocket motors is proposed. Actually, it is very difficult to have detailed information about the performance of the rocket motors beforehand because it is quite sensitive to combustion environments. To overcome this problem, we have developed an algorithm for generating in-flight prediction of rocket motor performance in realistic environments via a reference burnback profile and accelerations measured at a short time-interval just after launch. The performance is evaluated through a lot of flight test results.

The present paper discusses a need to develop a methodology of predicting the reliability of small thrust liquid rocket engines with a flow section made of composite materials under actual operating conditions for their successful practical use in the propulsion systems.

A system analysis program for the preliminary design of a liquid rocket engine is developed to analyze performance according to the input variables and the requirements needed. It contains sub-system analysis program modules that include the main thrust chamber, gasgenerator, turbo pumps, and turbine. The hydraulic system, which includes pipes, valves, and elbows, is considered to measure mass and pressure loss. The analysis process is developed to satisfy energy balance, mass-flow balance, and pressure balance. It also determines the design parameters with a higher specific impulse and thrust-to-weight ratio based on the genetic algorithm. The results of the program analysis are consistent with the reference of existing liquid rocket engines. The developed program is acceptable as a preliminary design process for a liquid rocket engine.

A systemic set of methods and devices designed to increase the maximum range of flight is considered. The main attention is paid to evaluation of efficiency for new technical solutions related to integrating the ramjet rocket engines into rocket vehicles with active and reactive launch principles. According to the results of studies, comparative assessments are presented for attainable values of the range increase based on the technical solutions considered.

Abstact This paper gives an overview of the Amelia small rocket program, developed to build low cost, reusable CanSat launchers and enable the construction of sounding rockets in Poland in the near future. The CanSat payload standard requires experiments to be containable in soda cans of 350 ml volume with a launch mass of up to 500 g. The paper covers early flown rocket designs as well as the recently tested two-stage Amelia 2 vehicle. A single stage, larger CanSat launcher is also presented. All of the vehicles are propelled by reusable composite solid rocket motors that will make it possible for small scientific experiments to be sent to altitudes of up to 10 km. One of the main goals is to develop rockets that are fully recoverable and reusable, thereby reducing operational costs. The sounding rocket that is viewed as a possible outcome of the presented research is also discussed. The work on the first rotating detonating engine flying demonstrator is also covered. This paper contains descriptions of the main subsystems of the rocket. Key technical facts and the vehicle visualizations are presented. The design methodology and calculation methods are also given. The paper contains a description of the technical equipment and background developed during the program, including propellant and motor production facilities and the launch pad. The results of the program, focused on using innovative propulsion technologies, are presented. This work is based on results of the joint project of the Space Technology Department of the Institute of Aviation and the Division of Aircraft Engines of the Institute of Heat Engineering of Warsaw University of Technology (WUT) with a key role played by members of the Students׳ Rocketry Group of the WUT Students׳ Space Association.

This paper describes the designing process of rocket engine with thrust 25N for system of orientation and starting of spacecraft instead the prototype with burnout chamber defect that was not fixed during the development to completely exclude the burnout possibility. The new version of the engine has been designed and additional research with CFD-modeling has been performed to completely exclude burnout possibility.

The inherent safety of hybrid rocket propulsion offers some unique advantages com- pared to solid and liquid propellant rocket engines. This makes it especially attractive for space tourism, Micro-launcher and hands-on experiments in the education of students. On November 8th, 2016 at 10:30 a.m. the hybrid sounding rocket HEROS 3 was launched from the ESRANGE Space Center to an apogee altitude of 32,300m (106,000 ft). This set a new altitude record for European student and amateur rocketry and a world altitude record for hybrid rockets built by students. The 7.5m long rocket was using Nitrous Oxide (N2O) and a Paraffin-based fuel to produce 10,000N of thrust. The dry mass of the rocket was only 75 kg thanks to a carbon fibre structure for the most part. The rocket performed the record breaking flight at perfect weather and visibility conditions, reaching a maximum airspeed of 720 m/s and Mach 2.3. The rocket performed a soft landing with two parachutes and can be reused. Flight data and engine performance data are published and analyzed. The flight data shows excellent stability of the rocket. Engine performance data proves very high efficiency and stable combustion as in the ground tests. The subsystem design and verification before the launch is reported. Engine and flight trajectory simulations show very good agreements with the flight data. Furthermore, the overall project, the rocket design, the subsystems as well as the launch campaign are presented here in detail.

The different types of aerospace propulsion engines are quantitatively described and the basic performance attributes of each by applying basic integral conservation equations. When work but no heat was added to the flow processed by the engine we had the case of the propeller, which turned out to be the most efficient propulsion device. Though efficient, the propeller was speed-limited because for a given power input, the thrust produced varied approximately inversely to the flight speed. When only heat is added but no net work is done on the fluid passing through engine, we have the case of the turbojet, ramjet, scramjet, or pulsejet engines, all of which are less efficient than the propeller, but capable of speeds much higher than those achievable with a propeller. A special case of adding only heat is the rocket, which takes in no ambient fluid but instead carries onboard all the propellants. The thrust of a rocket is independent of flight speed and because all the fluid it uses is carried onboard, the rocket is able to operate in the vacuum of space. By adding both work and heat to the fluid we arrive at the case of the turbofan which derives benefits of both the propeller and the turbojet. The ratio of the mass flow of the work-added part to the heat-added part is called the bypass ratio. The turbofan enjoys the lower specific fuel consumption provided by the fan section of the engine and the high-speed capability of the turbojet section of the engine. After illustrating the different engine types which use the principle of jet propulsion, attention was focused on the force field generated by such engines and the means by which the fundamental performance characteristics of the engines can be calculated.

A 70 microphone, 10×10 ft, microphone phased array was built for use in the harsh environment of rocket launches. The array was set up at NASA Wallops launch pad 0A during a static-test firing of Orbital Sciences’ Antares engines and again during the first launch of the Antares vehicle. It was placed 400 ft away from the pad and was hoisted on a scissor lift 40 ft above ground. The data sets provided unprecedented insight into rocket noise sources. The duct exit was found to be the primary source during the static-test firing; the large amount of water injected beneath the nozzle exit and inside the plume duct quenched all other sources. The maps of the noise sources during launch were found to be time dependent. As the engines came to full power and became louder, the primary source switched from the duct inlet to the duct exit. Further elevation of the vehicle caused spilling of the hot plume, resulting in a distributed noise map covering most of the pad. As the entire plume emerged from the duct, and ...

One of the main sources of concern within the design and manufacturing of rocket engines is the starting device, due to the relatively low reliability of the ignition of the main combustion process in these engines. Not only is necessary for the igniter to produce a sound ignition, but the starting pressure transient should manifest a tightly controlled build-up, in a very definite time interval. A specific research and development work was done by the authors in order to reliably assess the requirements for the ignition device of the newly developed compound, solid-liquid experimental rocket engine MEC-80 of the ADDA team. Computer predictions and scaled experiments on the constant volume combustion into a modified calorimeter were performed with notable results that led to the optimal design of the ignition device of the motor. The test results are presented with emphasize on the predictability of the ignition delay.

The flow rate control of a cavitating venturi has been investigated with downstream pressure variation. A set of cavitating venturies for a liquid rocket engine thrust chamber firing test facility have been designed and manufactured. The flow characteristics of the cavitating venturies have been analyzed by experimental and computational methods. Results showed that constant mass flow rate condition was established by the cavitation inside the venturi. However, upstream pressure less than the actual design pressure of the cavitating venturi could not supply a constant flow rate.

Hybrid rockets provide compelling features for use in atmospheric and space rocket propulsion. One of the prominent applications of hybrid rockets which foster on its characteristics is the propulsion of micro air launch vehicles. In this paper, a set of design options of a hybrid rocket motor is evaluated for propulsion of micro air launch vehicles. In order to evaluate the various design options of a hybrid rocket, we developed design and performance simulation codes. A simulation code is based on a legacy interior ballistic model. MATLAB® environment was used to develop the design and performance analysis codes and to visualize the temporal variation of performance characteristics and grain geometry during burning. We employ the developed codes to assess the replacement of solid rocket motors which are typically used in Air Launch Vehicles by hybrid rocket motors. A typical Micro Air Launch Vehicle mission to launch a 20-kg payload into a 400-km circular polar orbit is assumed. The results show that a hybrid rocket is a suitable candidate for micro air launch vehicles. The performance is improved in terms of specific impulse and thrust with smaller size in the same mission. Several design parameters of hybrid rocket motors were also evaluated and analyzed, including different fuel port geometry, type of fuels and oxidizers, number of ports, nozzle design and initial mass flux. These design parameters bring a significant effect on hybrid rocket performance and size.

Over the last decade, interest has been rekindled on hydrogen peroxide as a rocket propellant. As a result, 1000-N-class H2O2/hydrocarbon bipropellant rocket systems are in a stage requiring a serious exploration for application. Application to orbit-raising maneuvers is proposed here as the most appropriate use of the 1000-N-class thrusters. In this study, optimal design specifications of a 90% H2O2/kerosene rocket system for an apogee kick motor were derived. The optimal design specifications are configured with a thrust of 742 N, a chamber pressure of 11.7 bar, a mixture ratio of 7.36, and a nozzle expansion ratio of 580. It was found that, in the case of the latest Korean geosynchronous satellite, replacing the current apogee kick motor with the H2O2/kerosene rocket system would result in a mass increase of less than 0.5% of the total mass.

The flame deflector of a rocket engine test facility needs to be protected from the high thermal loads resulting from the impingement of high velocity and high temperature exhaust flame on it during the hot test. The flame deflector is cooled by water injected from holes in the deflector plate. In this paper ANSYS Fluent was used to analyze the flow over the flame deflector. The engine plume structure and its impingement characteristics on the flame deflector were determined in the absence and presence of water injection. The water injection pattern was optimized for the better cooling of the deflector plate. SST k-ω model was used for the turbulence modelling. Discrete Phase Model was used to simulate water injection from holes provided in the impingement plate.

This paper describes a test bench for rocket engines for educational purposes and small research launchers. Students can develop their own propulsion systems and test them under realistic process conditions. The test bench is equipped for solid combustion as well as for hybrid engines with solid propellant and gaseous oxidizer.

Before any flight test of a nuclear engine can take place, considerable testing of the engine will have to be performed both at the component and at the engine level. During the days of the NERVA program in the United States, open air testing of the engine was performed with little regard to the containment of fission products in the engine exhaust. Even though the contamination of the environment surrounding the NERVA test area was minimal, in today's regulatory environment, it is unlikely that such testing would be allowed again. As a result, the testing of nuclear rocket engines will probably require a variety of test facilities ranging from small nonnuclear facilities for component testing to very large facilities to test the entire engine assembly for long durations with total engine exhaust containment.

An air-breathing rocket engine inhales oxygen from the air for about half the flight, so it doesn't have to store the gas onboard. So at take-off, an air-breathing rocket weighs much less than a conventional rocket, which carries all of its fuel and oxygen onboard. Air breathing rockets, combine the performance characteristics of both rocket and ramjet engines. An air-breathing engine gets its initial take-off power from specially designed rockets, called air-augmented rockets, that boost performance about 15 percent over conventional rockets. When the vehicle's velocity reaches twice the speed of sound, the rockets are turned off and the engine relies totally on oxygen in the atmosphere to burn the hydrogen fuel. Once the vehicle's speed increases to about 10 times the speed of sound, the engine converts to a conventional rocket-powered system to propel the vehicle into orbit. And therefore reducing a vehicle's weight decreases cost significantly. And since an air breathing engine cannot get required initial take off thrust, various launch types like air augmented rockets, horizontal launch mode courtesy hybrid engine, magnetic levitation launch systems are used for initial thrust requirements ,thus reducing fuel emissions and increases net efficiency of rockets. Hence air breathing engines can be implemented to address energy considerations and reduce costs.

A design method for liquid-propellant rocket engines is proposed, on the basis of models adjusted in response to test results. This method increases the reliability of the results and reduces ground testing of the motors. The effectiveness of the method is confirmed by ground testing and flight experience with the motors developed by Salyut Design Bureau at Krunichev State Space-Research Center.

Because of the high-cost of the spaceflight launching, the reliability of the payload must be ensured in the high level acoustic energy produced by the first stage rocket engine. It’s the base for ensuring the normal work of the instruments during the lift-off. In this paper, the exhaust noise and the specific sound field produced by the liquid rocket engine on the ground-test system were measured, and the statistical energy analysis model was built to predict the noise environment of the rocket engine. In order to obtain the noise spectrum loaded on the fairing and the instrument cabin during the lift-off, two measurements were made and some measuring locations were set for each test. During the testing, the sound pressure level around the engine spout and the corresponding heights of the payload cabin were measured. By comparing the data derived from different measuring locations for the same measurement and the different measurements for the same location, the characteristics of the noise field was obtained. Results showed that the measured data agreed well with the predicted data used AutoSEA2 software.

This paper analyses the takeoff process of a small scaled UAV (unmanned aerial vehicle) with a single rocket booster. Because the thrust provided by the rocket booster is 10 times as large as the thrust provided by the engines, the effects caused by the boosting rocket on total mass, compound centre of gravity and inertia can not be neglected and are all considered. The inertia of the boosting rocket is calculated by the means of finite element method. Based on the analysis, a nonlinear dynamic model of the UAV is built. Several simulations with different takeoff parameters are conducted to test the takeoff performance. By analyzing simulation results, the acceptable range of boosting angle is investigated.

The method of analyzing the small-size rocket liftoff is presented. Under consideration is a system that consists of a rocket with an operating rocket engine, a launch tube, and an operating gas generator mounted in the tube under the rocket. The liftoff parameters including the rocket motion velocity inside the tube and the loads applied are calculated. The calculated and experimental data were compared and their adequate agreement is presented with respect both dynamics of rocket impulse-reaction liftoff from the launch tube and the launch velocity.

Integrated Systems Health Management (ISHM) consists of processes managing erroneous conditions that systems may encounter during their operational life by either designing out failures early on or defending and mitigating any possible failures. A successful implementation of ISHM consists of the following four components: data sensors, computations, data sinks, and visualization modules. In this paper, we explore the use of virtual reality (VR) platforms as a candidate for developing ISHM visualization modules. VR allows for a complete and spatially accurate 3-D model of a system to be displayed in real time. It provides a medium for improved data assimilation and analysis through its core tenants of immersion, interaction, and navigation. Furthermore, VR allows for integrating graphical, functional, and measurement data in the same platform-providing for the development of subsequent risk-analysis modules. The research objectives of this paper are focused on creating a detailed visual model of a multisensor rocket engine test facility inside a VR platform and demonstrating the capability of the VR platform in integrating graphical, measurement, and health data in an immersive, navigable, and interactive manner. A human-based performance evaluation of the VR platform is also presented. These research objectives are addressed using an example of a multisensor rocket-engine portable test stand at the National Aeronautics and Space Administration (NASA) Stennis Space Center's E-3 test facility.

A program that can simulate gas generator after burning liquid rocket engines was developed along with presenting the characteristics of the engines. The program was written in Matlab and used GUI interface so that many users can use it without any difficulties. The results of the program was compared with the real engine which was developed by the LRE advanced country. Most of the parameters concurred within 1% error expect for the pressure at the turbopump. The reasons of the large differences were supposed that pressure decreases at the schematics were smaller than that of the real engines.

Nuclear thermal propulsion (NTP) has been recognized as an enabling technology for missions to Mars and beyond. However, one of the key challenges of developing a nuclear thermal rocket is conducting verification and development tests on the ground. A number of ground test options are presented, with the Sub-surface Active Filtration of Exhaust (SAFE) method identified as a preferred path forward for the NTP program. The SAFE concept utilizes the natural soil characteristics present at the Nevada National Security Site to provide a natural filter for nuclear rocket exhaust during ground testing. A validation method of the SAFE concept is presented, utilizing a non-nuclear sub-scale hydrogen/oxygen rocket seeded with detectible radioisotopes. Additionally, some alternative ground test concepts, based upon the SAFE concept, are presented. Finally, an overview of the ongoing discussions of developing a ground test campaign are presented.

High efficiency rocket propulsion systems are essential for humanity to venture beyond the moon. Nuclear Thermal Propulsion (NTP) is a promising alternative to conventional chemical rockets with relatively high thrust and twice the efficiency of highest performing chemical propellant engines. NTP utilizes the coolant of a nuclear reactor to produce propulsive thrust. An NTP engine produces thrust by flowing hydrogen through a nuclear reactor to cool the reactor, heating the hydrogen and expelling it through a rocket nozzle. The hot gaseous hydrogen is nominally expected to be free of radioactive byproducts from the nuclear reactor; however, it has the potential to be contaminated due to off-nominal engine reactor performance. NTP ground testing is more difficult than chemical engine testing since current environmental regulations do not allow/permit open air testing of NTP as was done in the 1960's and 1970's for the Rover/NERVA program. A new and innovative approach to rocket engine ground test is required to mitigate the unique health and safety risks associated with the potential entrainment of radioactive waste from the NTP engine reactor core into the engine exhaust. Several studies have been conducted since the ROVER/NERVA program in the 1970's investigating NTP engine ground test options to understand the technical feasibility, identify technical challenges and associated risks and provide rough order of magnitude cost estimates for facility development and test operations. The options can be divided into two distinct schemes; (1) real-time filtering of the engine exhaust and its release to the environment or (2) capture and storage of engine exhaust for subsequent processing.

NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow.

The presentation describes the function of rockets to launch the space shuttle into the space, different processes involved in rocket manufacturing, fuel mixing process including ingredients of the fuel, casting, static test, igniter installation, shipment of the segments to Kennedy Space Center (KSC), assembly at KSC, etc. It also involves space Orbiter installation, launch preparation, launch, motor separation from the space shuttle, retrieving the rockets from sea and return for inspection, cleaning and refurbishment. Each rocket weighs about 1,100,000 pounds (of which 1,000,000 pounds is fuel). Both the motors burn for about 2 minutes (burning about 2,000,000 pounds of fuel) and generating about 6,000,000 pounds of thrust). The total length of the rocket assembly (rocket booster) is about 149 feet.

The Nimbus Project team, made of young engineers and university students, has developed a sounding rocket with the main aim to engineer a reusable, cost-effective test bench. This makes it possible to field test many aerospace technologies as guidance or reentry systems, enable airborne atmospheric measurements and high altitude students' experiments. The modular architecture allows the use of different payloads, composite material structures increase lightness while off-the-shelf components reduce costs; ease of transportation and usability qualify it as an extremely versatile platform. The system is scalable to carry heavier payloads and to perform scientific tests at higher altitudes. The technology demonstrator is going to reach an altitude of 1000m with cost-effective technology and components. Demonstrator's subsystems (rocket motor, main frame, chute deploy system, etc) have been both ground and flight tested and have shown an optimal accordance between theoretical model and testing. The demonstrator launch took place on 29th March 2015.

Liquid rocket engines for launch vehicles and space crafts as well as their subsystems need to be verified and qualified during hot-runs. A high test cadence combined with a flexible test team helps to reduce the cost for test verification during development/qualification as well as during acceptance testing for production. Test facility intelligence allows to test subsystems in the same manner as during complete engine system tests and will therefore reduce development time and cost. This paper gives an overview of the maturing of test engineering know how for rocket engine test stands as well as high altitude test stands for small propulsion thrusters at EADS-ST in Ottobrunn and Lampoldshausen and is split into two parts: • Part 1 gives a historical overview of the EADS-ST test stands at Ottobrunn and Lampoldshausen since the beginning of Rocket propulsion activities in the 1960s. • Part 2 gives an overview of the actual test capabilities and the test engineering know-how for test stand construction/adaptation and their use during running programs. Examples of actual realised facility concepts are given to demonstrate cost saving potential for test programs in both cases for development/qualification issues as well as for production purposes.

Combustion test facility for liquid rocket engine using kerosene and liquid oxygen has been developed for the purpose of cooling and performance study. Test engine of thrust under 1.0 KN can be evaluated, and the real combustion test ensures a good operation of the combustion test facility. Combustion test facility will be modified to supply natural gas and liquefied natural gas as fuel and to give a regenerative cooling test.

Performance figures are presented for reusable, winged rocket stages launched from several large transport aircraft, including the Boeing 747, the Russian Anotonov An-226, a large supersonic aircraft comparable to the XB-70 aircraft, which achieved Mach 3.1 flight with conventional turbojet propulsion in 1964, and a Mach 6 conceptual aircraft based on the MA145-XAB ramjet demonstrated in 1968. Advantages of air launch include a reduced ascent-to-orbit delta-velocity, reduced drag, the capability to launch at any latitude, and simplified abort modes. Launch from the XB-70 flight condition would allow a 22% reduction in delta-velocity required of the rocket and a significant reduction in drag incurred by rocket flight beginning at an atmospheric density 1/20th that of sea level. Launch from a Mach 6, 85,000-ft altitude condition would allow a 33% reduction in delta-velocity; ar eusable rocket of 262,000 lb launched at this condition could deliver a 23,500-lb payload to orbit utilizing a 523-s vacuum specific impulse advanced rocket engine. Fluorine/lithium‐hydrogen engines achieved 523 s in U.S. Air Force and NASA development programs of the 1960s and 1970s. Comparisons to ground-launched, all-rocket vehicles delivering equivalent payloads to orbit are presented. Fluorine propellant reactivity and engine development history are also discussed.

The objective of the 2013 Wisconsin Space Grant Consortium collegiate rocket competition was to construct a rocket to achieve an apogee of 3,000 feet and recover the rocket safely in flyable condition The rocket was designed to reach an apogee of 3000 feet. The main feature of the rocket includes passively controlling the drag to control the altitude reached. This is done by initial design of the airframe and fins. The other feature is deploying a drogue parachute on the ascent of the rocket in order to stop the rocket at 3000 feet. The rocket was designed to achieve an apogee greater than 3000 feet so that the deployment of a drogue parachute can stop the rocket at 3000 feet. Fiberglass tubing makes up the rocket due to its high strength and durability. Extra support was designed into the rocket to ensure that zippering does not occur. The rocket design included resulting in stable rocket based on the center of pressure and center of gravity. The recovery system of the rocket involves a dual deployment system. A drogue parachute is to be deployed at 2900 feet to stop the rocket at 3000 feet and control the rockets descent until the main parachute is deployed. The main parachute is then deployed at 700 feet and controls the rocket descent until landing on the ground. There is also a second set of ejection charges that will deploy the drogue parachute at apogee and the main parachute at 500 feet if the main charges do not deploy them. These ejection charges are fired by the MARSA4 and PerfectFlite StratoLogger altimeters. There is also a motor ejection charge backup based on a time delay to ensure the deployment of a parachute.

資料番号: SA1140113000

A fully reusable rocket vehicle is proposed to demonstrate good operability characteristics both on the ground and in e ight. For achieving technical readiness for future space transportation systems, design considerations not only for higher-performance-related issues but also those for good operability are needed. The proposed vehicle is to be used as a sounding rocket and has the capabilities of ballistic e ight, returning to the launch site, and landing vertically, making use of clustered liquid-hydrogen rocket engines. Before the development of this type of reusable rocket was initiated, a small test vehicle with a liquid-hydrogen rocket engine was built and e ight tested. A demonstration of vertical landing and exerciseof turnaround operation forrepeated e ightsare the major objectivesof the test vehicle. Two e ightswere performed in succession, and the e ight-test operation provided many valuable experiences for designing the fully reusable rocket vehicle.

Whereas NASA’ s current next-generation reusable launch vehicle research has largely focused on all-rocket single-stage-to-orbit vehicles, some attention is being given to advanced propulsion concepts such as rocket-based combined-cycle suitable for next-generation-and-a-half vehicles. Rocket-based combined-cycle engines combine rocket and airbreathing elements into a single integrated engine capable of multimode operation. Two options for e ight testing early versions of a rocket-based combined-cycle ejector scramjet engine have been investigated. The e rst option uses a single, subscale rocket-based combined-cycle engine module mounted to the X-34 air-launched technologytestbedforcaptiveteste ightsuptoaboutMach6.4.Thesecondoptioncombinesrocket-basedcombinedcycle engine e ight testing and envelope expansion with the simultaneous development of a new two-stage-to-orbit operational vehicle in the small Bantam payload class (220 lb) dubbed the W vehicle. To enable early testing and complement dual-use missile applications, both proposed e ight test engines evaluated use Earth storable propellants fortheirrocket primaries and hydrocarbon fuel fortheirairbreathing modes. Research resultsinclude preliminary sizing and performance data for each concept.

The effectiveness of a vapor-deposited Cu–Cr coating in protecting a NARloy-Z (Cu–3 wt.%Ag–0.5 wt.%Zr) substrate was investigated in a hydrogen/oxygen rocket engine test facility. The rocket facility generated high-temperature, high-pressure flows of combustion gas. The Cu–Cr-coated NARloy-Z article was actively cooled by cryogenic gaseous hydrogen so that the coating surface temperature was maintained below 650°C. The coating provided good protection for the NARloy-Z substrate under both hydrogen- and oxygen-rich hot gas environments. In erosion tests (hydrogen rich), the coating remained as a metallic mixture of Cu and Cr phases. In the erosion/oxidation tests (oxygen rich), the coating surface formed a protective Cr2O3 scale under a thin layer of Cu-oxides. The ability of the coating to withstand repetitive thermal shocks from the hot gas impingement without cracking was attributed to the close match of thermal expansion properties between the Cu–Cr coating and the NARloy-Z substrate.