Test results of the air turbo ramjet for a future space plane

The Institute of Space and Astronautical Science (ISAS) has been engaged in the development study on the Air Turbo Ramjet (ATR) engine since 1986 in cooperation with the Ishikawajima Harima Heavy Industries Co. Ltd (IHI). The ATR is one of the most preferable candidates for the propulsion system of a future space plane. Our ATR engine is a combined cycle air breathing propulsion system which consists of the turbojet and the fan boosted ramjet using the liquid hydrogen as a fuel. This engine system was named “ATREX” after employing the expander cycle. The ATREX is energized by thermal energy extracted regeneratively in both the pre-cooler installed in the air intake and the heat exchanger in combustion chamber. The ATREX works in the flight condition from sea level static up to Mach 6 at 35 km altitude. The ATREX employs the tip turbine configuration for compactness of turbo machinery. We are assessing the feasibility of the ATREX system by the sea level static tests using the 1/4-scale model (ATREX-500) with a fan inlet diameter of 300 mm and overall length of 2120 mm. In 1990, the ATREX-500 engine was tested in a sea level static condition to verify the performance characteristics of the turbo machinery and the combustor. In September of 1991, the heat exchanger was installed in the combustion chamber and tested independently from the turbo system. In November of 1991, the heat exchanger was coupled with the turbo system and tested to verify the overall system of the ATREX. In this paper are presented the test results of the ATREX-500 engine tested in the sea level static condition.     

Research on hypersonic aircraft using pre-cooled turbojet engines

Systems analysis of a Mach 5 class hypersonic aircraft is performed. The aircraft can fly across the Pacific Ocean in 2 h. A multidisciplinary optimization program for aerodynamics, structure, propulsion, and trajectory is used in the analysis. The result of each element model is improved using higher accuracy analysis tools. The aerodynamic performance of the hypersonic aircraft is examined through hypersonic wind tunnel tests. A thermal management system based on the data of the wind tunnel tests is proposed. A pre-cooled turbojet engine is adopted as the propulsion system for the hypersonic aircraft. The engine can be operated continuously from take-off to Mach 5. This engine uses a pre-cooling cycle using cryogenic liquid hydrogen. The high temperature inlet air of hypersonic flight would be cooled by the same liquid hydrogen used as fuel. The engine is tested under sea level static conditions. The engine is installed on a flight test vehicle. Both liquid hydrogen fuel and gaseous hydrogen fuel are supplied to the engine from a tank and cylinders installed within the vehicle. The designed operation of major components of the engine is confirmed. A large amount of liquid hydrogen is supplied to the pre-cooler in order to make its performance sufficient for Mach 5 flight. Thus, fuel rich combustion is adopted at the afterburner. The experiments are carried out under the conditions that the engine is mounted upon an experimental airframe with both set up either horizontally or vertically. As a result, the operating procedure of the pre-cooled turbojet engine is demonstrated.     

Results of the parabolic flight tests of the rapunzel deployer

The tether assisted re-entry of small payloads is a highly interesting tool for space transportation especially for the return of small payloads from Space Station ISSA. The small tether mission Rapunzel was initiated in 1991 by the Institute of Astronautics, TU München and the Kayser-Threde Company, to design a low cost and feasible tether experiment for the verification of the tether assisted re-entry. Together with the Samara State Aerospace University, Russia, a mission concept on a Russian Resurs or Photon capsule was developed. Based on this mission a deployer has been designed, mainly based on technology of the textile industry, which insures high reliability at low cost. Recently a similar configuration is being discussed for the ESA-TSE mission.The main work during the recent time was the development and test of the breadboard model of the deployer system. After successfully completing initial ground tests with the deployer, further tests during the ESA Parabolic Flight campaign in November 1995 were conducted. After a short introduction of the overall mission scenario, the planned configuration in orbit, this paper will present the results of the microgravity test campaign onboard the KC-135 aircraft and compare them with the ground test. The deployer showed a good performance during all tests, including ejection of the end-mass, deployment, and braking. Problems that occurred during the tests will be discussed, and solutions for the detected flaws and the results of the redesign now in progress will be presented. These verifications have shown the feasibility of the concept and will lay the base for the planned development of the flight model of the deployer.     

发动机喷流对飞行器飞行姿态影响的研究

针对某型导弹的实际飞行结果，在对立了合理的发动机推力模型及喷流模型的基础上，采用多学科的一体化仿真方法，重点研究了发动机喷流干扰对该型导弹飞行性能的影响。结果表明发动机喷流与发动机推力偏心及偏斜的共同作用将对飞行器飞行姿态和弹道产生显著影响，在飞行控制系统的仿真研究与参数整定中应综合考虑这些因素。     

变均方根随机振动?变加速度离心复合试验

为在地面试验中更真实地模拟再入飞行过载、振动力学环境的时变特性和复合特性,以再入飞行器子系统为对象,开展变均方根随机振动?变加速度离心复合试验研究。设定试验条件时,对预示弹道过载数据进行平滑和重采样,得到变加速度离心试验条件;对实测振动数据频谱进行归一化和频谱加权融合,得到变均方根随机振动试验条件。复合试验结果表明,给出的试验条件控制可行,实现了再入飞行振动历程和过载历程的同时模拟。     

MHD-Arc-Ramjet联合循环与AJAX间的性能比较

为了提高冲压发动机在更高速度区内的性能，扩展发动机的运行速度范围，需要在并行于发动机的燃烧室处附加一套能量旁路系统。采用热力学方法针对两种带有能量旁路的冲压发动机分别在理想和实际情况下进行了性能分析和比较，结果表明：理想条件下MAX的单位推力要高于MHD-Arc—Ramjet联合循环；考虑到后者可以在更高的温度下进行能量注入，其在单位推力方面是具有优势的。实际条件下在很宽的飞行Ma范围内MHD—Arc-Ramjet联合循环的单位推力要高于AJAX。     

DTFT-1: Analysis of the first USV flight test

The first dropped transonic flight test (DTFT) of the USV Program, performed with Castor, the first of the two spacecrafts developed within the USV Program, was performed on Saturday 24th February 2007, from Tortolì Airport in Sardinia.At 8:30 a.m. the 340 000 m³ stratospheric balloon lifted off from the East coast of Sardinia, bringing the flying test bed (FTB) up to 20.2 km before release within the isolated sea polygon controlled by Italian air force test range in Salto di Quirra (PISQ). The mission ended at 10:30 a.m. with the splash-down of the space vehicle.The flight itself was very good, with a nose-up manoeuvre under transonic conditions, reaching a maximum Mach as high as 1.08. The mission target was completely achieved as some 2 million measures were taken related to flight data, housekeeping, as well as 500 aerodynamic and structural experimental sensors. Unfortunately, the vehicle has been damaged more than expected during splash-down.Many national and international institutions and industries contributed to the mission carrying out, under the supervision and technical guide of CIRA: Italian Space Agency, Italian Air Force, Italian Navy, Italian Civil Aviation Authority, Italian Company for Air Navigation Services, Port Authorities, European Space Agency, Techno System Dev., Vitrociset, Carlo Gavazzi Space, Space Software Italia, Alcatel Alenia Space Italy, ISL-Altran Group.The paper reports the actual status of post-flight data analysis.     

新一代低温液体中型运载火箭（英文）

The maiden flight of LM-8 performed perfectly on December 22, 2020. The design concept of modularization, seriation and combination has been perfectly exhibited in LM-8. The four main technical innovations, including rapid integrated design based on modularization, engine thrust regulation, modal parameters acquisition technology based on numerical simulation, and flight load control, were verified during the maiden flight. LM-8 is now positioned to be the main force in China's medium launch vehicles for commercial launch. In the future, the mission adaptability of LM-8 will be improved to provide efficient and low-cost launch services. In addition, new technologies to allow repeated use and autonomous flight will be validated.     

分级燃烧循环发动机启动过程的变结构控制

针对分级燃烧循环液体火箭发动机启动过程的特点，提出了一种变结构控制系统。将启动过程的控制量分解为期望控制输入和随机反馈控制输入，后者由变结构控制来确定。选择发动机涡轮泵转速，预燃室和燃烧室压力作为跟踪状态变量构造线性切换函数，采用分段的等速趋近率实现滑动模态控制。这种变结构控制结构简单且具有较强的鲁棒性，使发动机启动过程的稳定性增强，仿真研究验证了变结构控制系统的有效性。     

基于差动气动增压装置的多次启动一体化动力系统特性

高性能、轻质量、小尺寸的上面级动力系统可以为整个任务节省更多的质量和空间,从而增加有效载荷的质量。提出了一种基于差动气动增压装置的多次启动姿轨控一体化动力系统,动力系统由主发动机系统、姿控发动机系统、一体化供应系统等组成。该动力系统的系统级试验验证结果表明:姿控发动机与主发动机可通过共用贮箱实现推进剂的一体化供应; 主...     

The analysis of measurements of quasi-stationary electrical fields in the near-surface zone of the <Emphasis Type="Italic">Mir</Emphasis> orbital complex

The quasi-stationary electrical fields have been measured in the near-surface zone of the Kvant module of the Mir orbital complex using vibration-type sensors of the Zond-Zaryad instrument. The analysis of measurements has shown that, depending on the spacecraft’s surface section shape (an octagon), four areas of the electrical fields configuration (and of the plasma component density) were revealed in the spacecraft vicinity: a ram with angular extension of about 120° (±60°), two side areas with angular extension of about 60° each, and a wake with angular extension of about 120° (±60°). The electrical field in the ram has a small absolute value. In the side areas, the electrical field can have either a negative or positive value, and it is smaller in magnitude than that in the ram. The electrical field in the wake has a three-petal structure (and an even more complicated structure is also possible): the negative value of the field (with angular extension of about 45°), the positive value of the field (with angular extension of about 30°), and, again, the negative value of the field (with angular extension of about 45°). Absolute value of the field in the wake is at least an order of magnitude higher than that in the ram and on the side surfaces.     

微重力科学与应用研究（上）

空间微重力环境的独特物理现象对航天器各系统均产生无法回避的影响，在型号研制中正确计入微重力效应是航天事业发展中的一个重要研究领域。在现阶段，航天器发动机再启动相关的微重力研究以及载人航天防火安全相关的微重力研究应当引起人们的高度重视。随航天事业的发展，微重力效应问题日益会严重起来（结构动力学、耦合动力学、多相流温控、低重星球场着陆撞击……），进行深入广泛的微重力效应研究对保证航天器飞行任务完成有着不应忽视的作用     

火箭发动机涡轮泵转速遥测数据实时处理方法

火箭飞行试验过程中,一般通过连续处理测量设备输出的脉冲周期计算发动机涡轮泵转速。受测控站接力跟踪及接收条件限制,遥测数据存在不连续、短时中断等现象,极易造成数据处理错误。通过研究脉冲参数在传输时的状态及状态转移,并对实时跟踪时出现的数据中断、异常值处理等情况进行分析,提出一种基于有穷自动机理论的处理模型,有效解决了测量数据的连续判断问题。     

发动机喷流干扰对底部热环境影响研究

发动机底部热环境的准确预示是小型运载火箭研制的关键环节。为提升对小型运载器底部热环境的认识,开展了发动机喷流干扰对底部对流加热影响的研究。首先,采用计算流体力学方法,开展了发动机喷流流动的数值计算研究,分析了飞行高度、发动机开关机、飞行攻角对底部对流加热的影响;然后,从流动机理出发,提出了一种降低底部对流加热的外形优化方法;最后,根据飞行试验测量结果,讨论了底部加热的主要来源。     

Benefits from incorporation of combined cycle propulsion

The X-33 program was initiated to develop a testbed for integrated RLV technologies that pave the way for a full scale development of a launch vehicle (Venture Star). Within the Nasa Future X Trailblazer program there is an Upgrade X-33 that focuses on materials and upgrades. The authors propose that the most significant gains can be realized by changing the propulsion cycle, not materials. The cycles examined are rocket cycles, with the combustion in the rocket motor. Specifically, these rocket cycles are: turbopump, topping, expander, air augmented, air augmented ram, LACE and deeply cooled. The vehicle size, volume, structural weight remain constant. The system and propellant tank weights vary with the propulsion system cycle. A reduction in dry weight, made possible by a reduced propellant tank volume, was converted into payload weight provided sufficient volume was made available by the propellant reduction. This analysis was extended to Venture Star for selected engine cycles. The results show that the X-33 test bed could carry a significant payload to LEO (10,000 Ib) and be a valuable test bed in developing a frequent flight to LEO capability. From X-33 published information the maximum speed is about 15,000 ft/sec. With a LACE rocket propulsion system Venture Star vehicle could be sized to a smaller vehicle with greater payload than the Venture Star baseline. Vehicle layout and characteristics were obtained from: http:// www.venturestar.com.     

Equivalence ratio and constriction effects on RBCC thrust augmentation

A theoretical analysis of a variable area rocket based combined cycle engine with and without simultaneous mixing and combustion is presented. The flowfield is solved using a steady, quasi-one-dimensional, inviscid control volume formulation with combustion effects included via a generalized equilibrium calculation. Compression augmentation is shown to be sensitive to the equivalence ratio within the primary rocket chamber, where ejector section performance is greatest at both low and high equivalence ratios but near a minimum at stoichiometric conditions. The thrust generated by the RBCC engine compared to that generated by the same rocket in isolation can be increased by as much as 12% at constriction ratios of between 45% and 50%. Thrust augmentation is also shown to vary with equivalence ratio, where for a fixed geometry the maximum thrust is generated at equivalence ratios slightly below unity.     

The modular optelectronic multispectral scanner system for spaceborne remote sensing

A multispectral scanner system for spaceborne remote sensing of land and coastal/ocean features is under development for the German Ministry for Research and Technology. The system is based on the use of multilinear detector arrays for visible and infrared spectral bands.The electronically scanning image system MOMS (Modular Optoelectronical Multispectral Scanner) consists of individual spectral channel modules which can be grouped to dedicated mission tasks. Those dedicated tasks are land surface thematic mapping, sea or vegetation monitoring and in a stereo mode conventional photo interpretation and mapping.The basic performance data would allow up to 10,000 pixels per scan line, corresponding to about 20 m resolution at 200 km swath width out of observation satellite altitudes with narrower spectral bands than used on the current systems. High spectral resolution (up to 20 nm) is feasible at medium spatial resolution (～ 60 m).An experimental airborne scanner has been successfully flown in spring 1978. High-resolution modules development in the visible/NIR is under way and will be flight tested in early 1981.     

Human thermohomeostasis onboard "Mir" and in simulated microgravity studies.

Significant changes of thermogomeostatic parameters was obtained by thermotopometric method using the techniques simulate of microgravity effects: bed rest, pressurized isolation, suit immersion (SI). However, each of ground models made rectal temperature (T) trend downward. The autothermometric study (24 and 12 sessions, 2-13th and 6-174th flight days) was carried out onboard "Mir" by two flight engineers who had preliminary tested at SI (1-2 days). Studies of German investigators onboard "Mir" confirmed: rectal T must be higher in space flight as compared to the normal environment (n=4). Comparative studies suggest that microgravity is a key factor for the human body surface T raise and abolishment of the external/internal T-gradient. T-homeostasis was not really changing during missions and could be regarded as acute effect of microgravity. After delineation of changes in body surface T--by Carnot's thermodynamic law--rectal T raise should have been anticipated. Facts pointing to the excess entropy of human body must not be passed over.     

空间站空空支架天线热设计与仿真分析

为研究核心舱飞行姿态、空间外热流、核心舱发动机羽流参数以及天线外表面热控涂层对空间站空空支架天线温度的综合影响,验证天线被动热控设计的有效性,进行了2种低温工况和6种高温工况的热分析。结果显示:低温工况下,通信天线惯性飞行时的最低温度低于正向飞行时的;展开臂多层表面最低温度为-85 ℃,满足温控指标。高温工况下,通信天线惯性飞行时的温度高于正向飞行时的;轨控发动机的羽流热效应大于偏航发动机的。通信天线内外表面均喷涂ACR-1温控白漆,1倍轨控发动机羽流热流密度时,最高温度为123 ℃,可满足实际使用要求。     

适应液氧煤油发动机变推力工况的伺服机构设计方案（英文）

An electro-hydraulic actuator for the thrust vector control(TVC) of a throttlable kerolox rocket engine is introduced in this paper. The creative feature is an integrated hydraulic power drive unit, where a constant speed kerosene motor is used to draw high pressure kerosene from the engine and to drive a constant pressure variable displacement piston pump, acting as the power supply for the actuator. Its operational mechanism, to accommodate the varying pressure from the turbo-pump of a throttling engine, lies in a pressure-reducing flow regulator inserted at the motor inlet. Another key point is that the displacement of the motor is reasonably bigger than the pump so that a sufficiently wide range of pressures can be adapted. Modeling analysis and flight test results were well matched, which show the outstanding performance of this novel type actuator.