Space Transportation Vehicles Research Articles

Thrust performance of the rocket-based combined-cycle engine under ejector mode

The Rocket-Based Combined-Cycle (RBCC) engine integrates the preponderance of ramjet and rocket engines. It can perform excellently at a lower Mach number than a ramjet while consuming less fuel than the rocket. The higher specific impulse under lower flight Mach conditions guarantees the competitiveness and application prospect of an RBCC engine for reusable space transportation and hypersonic cruise vehicles. With a wide range of Mach numbers, the flow choking between primary and secondary streams in the inner flow passage of the engine becomes complicated. The flow choking not only affects the mass flow ratio between the air stream and rocket plume but also determines the thrust performance of the engine. However, the relationship between flow-choking states and thrust performance has not been revealed. This investigation aimed to provide a better design reference for the RBCC trajectory design and a basis for the RBCC engine geometry design so that the thrust performance under different flow choking states was studied. The findings indicate that the engine is more favorable in the supersonic regime status than others. At the premise of the lower total pressure air stream condition, the thrust of the RBCC engine is attributed to the rocket pressurization effect. On the contrary, its thrust is ascribed to the air stream when the air stream total pressure is higher. Besides, the thrust augmentation of the RBCC engine is essentially due to the larger entrainment ratio. There is a rocket mass flow rate range in different mixer diameters, named region B in this paper, in which the air mass flow rate can be boosted with the rocket mass flow rate and the specific impulse ratio is higher than the rocket specific impulse. Significantly, the RBCC design state should be maintained in region B to meet the operational requirements of engines under the precondition of an insufficient flight Mach number.

Paste Propulsion of Space Transportation Vehicles for Orbital Transfer Applications

Paste Propulsion of Space Transportation Vehicles for Orbital Transfer Applications

Thermal Investigation of a Purged Insulation System for a Reusable Cryogenic Tank

To advance reusable space transportation vehicles with cryogenic tanks for future space missions, a reliable thermal management system must be developed. To do so, the challenge of insulating the propellant while at the same time protecting the structure against reentry heat loads needs to be addressed. A possible insulation design including a structural fixation element for the thermal protection system was developed as part of the German Aerospace Center project AKIRA. As a key feature a purged gap was implemented between cryogenic and high-temperature insulation to reduce overall layer thickness. Thermal experiments simulating both steady-state refueling phase and transient reentry phase proved that the defined thermal boundary conditions of the reference configuration could be met. The insulation system design is outlined in brief and both the experimental setup and the test results are presented and discussed.

Hydrogen Leakage Detection of Reusable Rocket Experiment Vehicle with Hydrogen/Oxygen Concentration Ratio Sensor

In future space transportation system, concept of fault tolerant is important. Propellant leakage is one of fault modes that should be considered in the system. The space transportation vehicles could have flight environment with very little or without oxygen. Therefore, for the propellant leakage detection, gas sensors that can work without oxygen are needed. The authors have developed the gas sensor that can work without oxygen and of which output depends on the ratio of hydrogen to oxygen concentration. The present sensors were installed on a hydrogen-fueled reusable rocket experimental vehicle so that the characteristics of the sensor’s behavior was measured and evaluated though ground firing test operation of the rocket engines. Comparison of the present sensor’s behavior to that of reference sensors shows validity of the present sensor.

Research and development on inlets for rocket based combined cycle engines

Rocket based combined cycle (RBCC) engine is regarded as one of the most efficient propulsion systems for reusable space transportation vehicles and hypersonic cruise vehicles. It can perform excellently over a very broad range of Mach numbers and flight altitudes through multi-mode operations by effectively integrating high thrust-to-weight ratio of the rocket engine and high specific impulse of the airbreathing engine. The RBCC inlet, commonly operating in an extremely wide flight speed range covering static, subsonic, transonic, supersonic, and hypersonic, plays a vital role in the RBCC propulsion system and flight vehicles. This paper presents a comprehensive and systematical review of the research progress on the inlets for RBCC engines in different countries and different historical periods. A number of key technical and scientific issues specially concerning RBCC inlets are sorted out, including the wide range RBCC inlet design and optimization, inlet/embedded rocket/combustor matching and control, dynamic operation, etc. Several typical RBCC inlet applications and the corresponding research are discussed, including the Strutjet inlet, A5 inlet, GTX inlet, and the Astrox's inlet in the U.S., the RBCC inlet in JAXA, and a sidewall compression RBCC inlet in China. Finally, experience gained and some recommendations are summarized for future research of the RBCC inlet.

Forming cyclogram of energy-propulsion system for prospective inter-orbital space transportation vehicle with electric propulsion and liquid stages

Forming cyclogram of energy-propulsion system for prospective inter-orbital space transportation vehicle with electric propulsion and liquid stages

Experimental and Numerical Investigations of Slender Body Side Force with Asymmetric Protuberances

Space transportation vehicles have asymmetric protuberant devices on their surfaces, such as cable ducts. Such protuberances, typically arranged asymmetrically with respect to the vehicle axis, are known to cause asymmetric vortices and side force. In this study, to understand effects of side force and its associated flow fields, both wind-tunnel tests and numerical calculations for a slender body with an asymmetric protuberance were conducted at Mach 1.5. The results of computed aerodynamic coefficients are in good agreement with the experimental results and detailed flow structures are provided. In particular, the results revealed that side force was generated by two factors. It linearly increased as vortices detached from the body, and nonlinearly increased based on the effects of secondary vortex scale as the angle of attack increased. Additionally, the axial position and azimuthal angle (angle along the circumferential direction around the body axis) of the protuberance strongly influenced side force characteristics. First, the side force was significantly higher when the protuberance was installed in a forward axial position. Second, when the protuberance was installed on the leeward side of the slender body, the side force increased with the angle of attack. These results are not limited to the presented configuration but to other rocket designs.

Health structure monitoring for the design of an innovative UAS fixed wing through inverse finite element method (iFEM)

The structural health monitoring (SHM) plays an essential role in system health management applications for aeronautic and space transportation vehicles, manned and unmanned. The unmanned aircraft systems (UAS) are also extremely needed in various fields of interest, from military to civilian (search and rescue, environmental surveillance and monitoring, entertainment). This work presents an innovative UAS fixed-wing design and control through an inverse finite element method-based, which compute the full-field displacements reconstruction of a three-dimensional shell/plate deformations from experimentally measured surface strains. The full-field displacements are useful for the preliminary design and inspections of the UAS loads, caused by maneuvers or gusts. Goal of this paper was to validate the high accuracy predictions of deformations afforded due to the inverse finite element method (iFEM). Overall formulation was based on the minimization of a least-squares functional that uses and compares the strains extracted due to embedded sensors with the strains of linear, first order shear-deformation theory. The test article was a thin plate equipped with embedded sensors (strain gauge sensors) which permit to extract surface strains in real-time, used as input data for shape sensing. The plate was used to approximate an UAS wing box section, in further work analyzed.

Develop Global Safety Synergies for Long-Range Human Space Exploration, with Focus on Launch Systems

ABSTRACT The agencies participating in the International space Exploration Coordination Group (ISECG) demonstrate a natural synergy for defining and implementing meaningful steps towards advanced human spaceflight. Within the exploration mission themes of the Global Exploration Roadmap (GER) [1], crew transportation and operation capabilities are driven by key requirements based on human and robotic mission partnership. They require an acceptable level of safety associated with the risks inherent to live, travel and work in space. This paper pursues the goal to investigate the safety orientations to be targeted by the space agencies, and addresses the main decision factors for implementing global safety rules regarding the space transportation vehicles.

EFFECT OF GROUND PROXIMITY ON AERODYNAMIC FORCES AND MOMENTS OF REUSABLE LAUNCH VEHICLE AT SUBSONIC SPEED

This paper explores the effect of ground proximity on aerodynamics forces and moments acting on reusable launch vehicle at the time of its landing. As the name indicates the concept of reusability is attractive especially in case of Space Transportation and launch vehicles as it reduces the mission cost. The design of Reusable Launch vehicle is also very critical as the vehicle has to pass through various speed ranges from hypersonic to subsonic. RLV wing is generally low aspect ratio double delta wing which is very sensitive to ground effect due to the characteristic excessive vehicle attitudes associated with the landing and touchdown phases of flight.. In present study, a RLV model of 260 mm total length and reference area of 155mm2 had been modeled using CAD tools. It has double delta wing of wing span 144mm and mean chord of 76mm and vertical stabilizer of height 38mm. Effect of ground proximity on aerodynamic forces acting on RLV has been studied using commercial CFD software. During present analysis computational analysis has been done for different height. Height of RLV is varied from free flight condition to landing touchdown height and the angle of attack has been varied for 0 to 20. The complete study is done for lower subsonic speed at Reynolds number of 1.8*106 for fix ground and moving ground conditions separately.

Numerical study on aerodynamic heat of hypersonic flight

Accurate prediction of the shock wave has a significant effect on the development of space transportation vehicle or exploration missions. Taking Lobb sphere as the example, the aerodynamic heat of hypersonic flight in different Mach numbers is simulated by the finite volume method. Chemical reactions and non-equilibrium heat are taken into account in this paper, where convective flux of the space term adopts the Roe format, and discretization of the time term is achieved by backward Euler algorithm. The numerical results reveal that thick mesh can lead to accurate prediction, and the thickness of the shock wave decreases as grid number increases. Furthermore, most of kinetic energy converts into internal energy crossing the shock wave.

Numerical modeling of self-pressurization and pressure control by a thermodynamic vent system in a cryogenic tank

This paper presents a numerical model of a system-level test bed—the multipurpose hydrogen test bed (MHTB) using the Generalized Fluid System Simulation Program (GFSSP). MHTB is representative in size and shape of a space transportation vehicle liquid hydrogen propellant tank, and ground-based testing was performed at NASA Marshall Space Flight Center (MSFC) to generate data for cryogenic storage. GFSSP is a finite volume-based network flow analysis software developed at MSFC and used for thermofluid analysis of propulsion systems. GFSSP has been used to model the self-pressurization and ullage pressure control by the Thermodynamic Vent System (TVS). A TVS typically includes a Joule–Thompson (J–T) expansion device, a two-phase heat exchanger (HEX), and a mixing pump and liquid injector to extract thermal energy from the tank without significant loss of liquid propellant. For the MHTB tank, the HEX and liquid injector are combined into a vertical spray bar assembly. Two GFSSP models (Self-Pressurization and TVS) were separately developed and tested and then integrated to simulate the entire system. The Self-Pressurization model consists of multiple ullage nodes, a propellant node, and solid nodes; it computes the heat transfer through multilayer insulation blankets and calculates heat and mass transfer between the ullage and liquid propellant and the ullage and tank wall. A TVS model calculates the flow through a J–T valve, HEX, and spray and vent systems. Two models are integrated by exchanging data through User Subroutines of both models. Results of the integrated models have been compared with MHTB test data at a 50% fill level. Satisfactory comparison was observed between tests and numerical predictions.

Advanced Flight Management System for an Unmanned Reusable Space Vehicle

Sangam M, Sabatini R, Ramasamy S and Gardi A. (2013). Advanced flight management system for an unmanned reusable space vehicle. International Journal of Unmanned Systems Engineering. 1(3): 48-67. The innovative architecture of an advanced Flight Management System (FMS) for Unmanned Reusable Space Vehicle (URSV) applications is presented with the associated re-entry trajectory computation algorithm. The SL-12 unmanned space vehicle, developed by Cranfield University as a part of the 2012-2013 Aerospace Vehicle Design (AVD) Group Design Project (GDP) is used as the reference platform. The overall avionics architecture of the future space transportation vehicle is described. A detailed architecture is developed for the FMS and the core functions of such an FMS are described. A dedicated computation algorithm is presented for re-entry trajectory planning, which involves determination of the path of re-entry vehicle by means of angle of attack and bank angle modulation. Simulation case studies are performed in a realistic re-entry operational scenario resulting in the generation of efficient and feasible trajectories, without violating any of the defined constraints. © Marques Engineering Ltd.

Conceptual study of manned space transportation vehicle using laser thruster in combination with the H-II rocket

This paper describes the conceptual study of a Manned Space Transportation Vehicle (MSTV) using a laser thruster in combination with the H-II Rocket. By combining the use of a laser thruster and H-II Rocket, space trip to the International Space Station (ISS) or a round trip mission around the moon can be performed. Once MSTV with one crew achieves a circular orbit at an altitude of 200km around the earth (parking orbit) by use of H-II Rocket, MSTV will then put into a circular orbit into an altitude of 400km (ISS orbit) from 200km circular orbit by use of the laser thruster. H-II Rocket has the following launch capability with payloads for LEO (300km): 10t (H-II A Rocket), 16.5t (H-II B Rocket). Laser thruster using water propellant, power source for the laser, orbital transfer calculations (to ISS or the Moon) and other practical aspects are examined.

無人宇宙輸送船「こうのとり」―安全設計への取り組み―

無人宇宙輸送船「こうのとり」―安全設計への取り組み―

A reliability, maintainability, and safety model to support the assessment of space vehicles

PurposeThe focus of this paper is on reliability and availability design goals. It aims to provide top‐level estimates of the safety and maintainability of future spacecraft systems.Design/methodology/approachThe developed design tool uses basic reliability principles to estimate the probability of a safe mission and the need for repairs/replacement during ground processing, before launch and start of mission, based on the characteristics of the vehicle's main systems: the number of subsystems, the mean time to repair, and the per subsystem average reliability.FindingsA simple reliability, maintainability and safety model is developed to support the top‐level design process of future space transportation vehicles. It also describes how the developed design tool uses various sensitivity analysis functions to improve design decisions.Originality/valueThe goal of the developed tool is to provide engineers/vehicle developers during the early stages of design with a tool that demonstrates the effect on maintainability of improving component reliability and reducing the number of components.

極超音速衝撃波・境界層干渉におけるCrackの空力加熱率への影響

A small crack on body surface led to a tragic accident in 2003, which is the Columbia accident. During the shuttle's re-entry, high temperature gas penetrated crack on leading-edge of the left wing and melted the aluminum structure, finally the Columbia blew up. Since early times, there are many fundamental studies about simple cavity-flow formed on body surface in hypersonic speeds. However, an investigation of Shock/Boundary-Layer Interaction (SBLI) on crack has not been researched. For multistage space transportation vehicle such as TSTO, SBLI is an inevitable problem, and then SBLI on crack becomes a critical issue for TSTO development. In this study, the effects of crack, where SBLI occurs, were investigated for TSTO hypersonic speed (M∞ = 8.1). A square crack locates at SBLI point on the TSTO booster. Results show that a crack and its depth strongly effect on peak heat flux and aerodynamic interaction flow-field. In the cases of shallow crack (d/C ≤ 0.10), there exist two high heat flux regions on crack floor, which locates at a flow reattachment region and a back end wall of crack. In this case, a peak heat flux at flow reattachment region becomes about 2 times as large as the stagnation point heat flux, which value becomes larger compared with a peak heat flux in the case of No-Crack TSTO. While in the case of deep crack (d/C = 0.20), overall heat flux on crack floor decreases to below the stagnation point heat flux. These results provide useful data for a development of TSTO thermal protection system (TPS) such as thermal protection tile.

Trajectory Optimization and Conceptual Study of Small Test Vehicles for a Hypersonic Engine Using a High-Altitude Balloon

The Japan Aerospace Exploration Agency, JAXA, announced a long-term vision recently. In the vision, JAXA aims to develop hypersonic aircrafts. A pre-cooled turbojet engine has great potential as one of newly developed hypersonic airbreathing engines. We also expect the engine to be installed in space transportation vehicles in the future. For combustion test in the real flight conditions of the engines, JAXA has an experimental plan where a small test vehicle is released from a high-altitude balloon. This paper applies numerical analysis and optimization techniques to conceptual designs of the test vehicle in order to obtain the best configuration and trajectory for the flight test. The results show helpful knowledge for designing prototype vehicles.

B211 PIVによる再使用ロケット周りの流れの可視化

A vertical take-off and vertical landing rocket is one of the future space transportation vehicle expected as fully reusable system. In the landing phase of vertical lander, the vehicle is decelerated by the main engine thrust and lands softly to the ground site. Then its aerodynamic characteristics are affected by the interaction between the engine plume and the subsonic free-stream against the vehicle. In order to investigate the influence of such interaction, wind tunnel tests were conducted. The aerodynamic forces and surface pressure were measured by using scale model of the Reusable Vehicle Testing (RVT) which is a small vehicle built for flight tests in ISAS/JAXA. Flowfield around the vehicle model was visualized by using Particle Image Velocimetry (PIV) method. As a result, the drag force and pitching moment acting the vehicle were affected by the change of pressure distribution due to the jet/free-stream interaction.

高高度気球を用いた極超音速エンジン小型実験機に対する軌道最適化と概念検討

Japan Aerospace Exploration Agency, JAXA announced a long-term vision recently. In the vision, JAXA aims to develop hypersonic aircrafts. A pre-cooled turbojet engine has great potential as one of newly developed hypersonic air-breathing engines. We also expect the engine to be installed in space transportation vehicles in future. For combustion test in real flight condition of the engines, JAXA has an experimental plan with a small test vehicle falling from a high-altitude balloon. This paper applies numerical analysis and optimization techniques to conceptual designs of the test vehicle in order to obtain the best configuration and trajectory that can achieve the flight test. The results show helpful knowledge when we design prototype vehicles.