Energetic Compositions Application for the Reduction of the Environmental Pollution Because of Space Vehicle Launches

Technogeneous impact of rocket and space activities on the environment is one of the most actual problems of practical cosmonautics. This technogeneous impact is not only the pollution of near Earth space with space debris (worked-off stages of space launch vehicle (SLV)), but also the pollution of significant areas on the Earth surface with worked-off lower stages of SLV, which fall down after having accomplished their mission. In OmSTU and IPCP RAS it was suggested to apply different self-burning compositions, generating hot gases for the evaporation of the unused residues of liquid propellant in tanks of SLV. Then the mixture of the evaporated compounds together with the gaseous combustion products from gas-generating compositions is used as propellant mixture for the autonomous gas rocket engine. Such a solution would decrease considerably the level of the environment pollution and additionally it increases the energetic characteristics of SLV. For example, in the case of the second stage of SLV «Soyuz-2.1.v» it increases the total velocity by 5%. Also it is proposed to use firing the pyrotechnic compositions like (thermites) for the fairings heating up to the temperature when the fairing material can be ignited in air. It would reduce considerably the amount and the mass of the separating parts of SLV that fall to the Earth

Development of Solid Gas Generating Compositions to Ensure Non Explosiveness of Spent Orbital Stages of Liquid Rocket of Space Launch Vehicles

The choice is discussed of solid gas generating compositions for venting by hot combustion products a fuel tank of the spent orbital stage of a space launch vehicle with a main liquid rocket engine. Non explosiveness is achieved via eliminating the possibility of freezing the drainage system when products of gasification (vapours of a propellant component + the remains of a gas boost + the hot products of combustion of solid gas generating compositions) are discharged from the tank into surrounding space. There are imposed requirements, constraints, and criteria for selecting solid gas generating compositions. When considering tank with the residues of liquid oxygen belonging to orbital spent stage of the launch vehicle В«ZenithВ» the ways are shown how to ensure explosion safety, which on the basis of proposed approaches by selecting solid gas generating compositions (SGC) which generate oxygen and nitrogen. As a criterion of choice of SGC the total mass of the gasification system is adopted, which includes the SGC mass for gasification of liquid propellant residues, the mass of the gas generator and the mass of system to supply the combustion products of SGC into the tank. It is proposed use of residual heat in the condensed phase of the SGC combustion products to heat up the drainage system, which will increase the probability of a trouble-free operation of the drainage system

Active debris multi-removal mission concept based on hybrid propulsion

During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to about 7000 metric tons. Most of the cross-sectional area and mass (97% in low Earth orbit) is concentrated in about 4500 intact abandoned objects plus a further 1000 operational spacecraft. Analyses have shown that the most effective mitigation strategy should focus on the disposal of objects with larger cross-sectional area and mass from densely populated orbits. Recent NASA results have shown that the worldwide adoption of mitigation measures in conjunction with active yearly removal of approximately 0.2–0.5% of the abandoned objects would stabilize the debris population. Targets would have typical masses between 500 and 1000 kg in the case of spacecraft, and of more than 1000 kg for rocket upper stages. In the case of Cosmos-3M second stages, more than one object is located nearly in the same orbital plane. This provides the opportunity of multi-removal missions, more suitable for yearly removal rate and cost reduction needs. This paper deals with the feasibility study of a mission for the active removal of large abandoned objects in low Earth orbit. In particular, a mission is studied in which the removal of two Cosmos-3M second stages, that are numerous in low Earth orbit, is considered. The removal system relies on a Chaser spacecraft which performs rendezvous maneuvers with the two targets. The first Cosmos-3M stage is captured and an autonomous de-orbiting kit, carried by the Chaser, is attached to it. The de-orbiting kit includes a Hybrid Propulsion Module, which is remotely ignited to perform stage disposal and controlled reentry after Chaser separation. Then, the second Cosmos-3M stage is captured and, in this case, the primary propulsion system of the Chaser is used for the disposal of the mated configuration. Critical mission aspects and related technologies are investigated at a preliminary level. In particular, an innovative electro-adhesive system for target capture, a mechanical system for the hard docking with the target and a hybrid propulsion system suitable for rendezvous, de-orbiting and controlled reentry operations are analyzed. This is performed on the basis of a preliminary mission profile, in which suitable rendezvous and disposal strategies have been considered and investigated by numerical analysis. A preliminary system mass budget is also performed, showing that the Chaser overall mass is about 1350 kg, including a primary propulsion system of about 300 kg and a de-orbiting kit with a mass of about 200 kg. This system is suitable to be launched with VEGA, actually the cheapest European space launcher

Active space debris removal by a hybrid propulsion module

During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to a total tally of approximately 7000 metric tons. Now, most of the cross-sectional area and mass (97% in LEO) is concentrated in about 4600 intact objects, i.e. abandoned spacecraft and rocket bodies, plus a further 1000 operational spacecraft. Simulations and parametric analyses have shown that the most efficient and effective way to prevent the outbreak of a long-term exponential growth of the catalogued debris population would be to remove enough cross-sectional area and mass from densely populated orbits. In practice, according to the most recent NASA results, the active yearly removal of approximately 0.1% of the abandoned intact objects would be sufficient to stabilize the catalogued debris in low Earth orbit, together with the worldwide adoption of mitigation measures. The candidate targets for removal would have typical masses between 500 and 1000 kg, in the case of spacecraft, and of more than 1000 kg, in the case of rocket upper stages. Current data suggest that optimal active debris removal missions should be carried out in a few critical altitude-inclination bands. This paper deals with the feasibility study of a mission in which the debris is removed by using a hybrid propulsion module as propulsion unit. Specifically, the engine is transferred from a servicing platform to the debris target by a robotic arm so to perform a controlled disposal. Hybrid rocket technology for de-orbiting applications is considered a valuable option due to high specific impulse, intrinsic safety, thrust throttle ability, low environmental impact and reduced operating costs. Typically, in hybrid rockets a gaseous or liquid oxidizer is injected into the combustion chamber along the axial direction to burn a solid fuel. However, the use of tangential injection on a solid grain Pancake Geometry allows for more compact design of the propulsion unit. Only explorative tests were performed in the past on this rocket configuration, which appears to be suitable as de-orbiting system of new satellites as well as for direct application on large debris in the framework of a mission for debris removal. The paper describes some critical aspects of the mission with particular concern to the target selection, the hybrid propulsion module, the operations as well as the systems needed to rendezvous and dock with the target, and the disposal strateg

Active debris removal space mission concepts based on hybrid propulsion

During the last 40 years, the mass of the artificial objects in orbit increased quite steadily at the rate of about 145 metric tons annually, leading to a about 7000 metric tons. Most of the cross-sectional area and mass (97% in LEO) is concentrated in about 4500 intact abandoned objects plus a further 1000 operational spacecrafts. Analyses have shown that the most effective mitigation strategy should focus on the disposal of objects with larger cross-sectional area and mass from densely populated orbits. Recent NASA results have shown that worldwide adoption of mitigation measures in conjunction with active yearly removal of approximately 0.1% of the abandoned objects would stabilize the debris population. Targets would have typical masses between 500 and 1000 kg in the case of spacecraft, and of more than 1000 kg for rocket upper stages. In the case of Cosmos-3M 2nd stages, more than one object are located nearly in the same orbital plane. This provides the opportunity of multi-removal missions, more suitable for yearly removal rate and cost reduction needs. This paper identifies the requirements of two alternative space mission concepts for multiple active removal of large debris objects from LEO, as well as at investigating related technological and operational issues. The first concept relies on flying a multi-removal space platform carrying a number of Hybrid Engine Modules (HEMs) to dock with the pre-selected targets, according to a pre-fixed removal sequence. Each target is then de-orbited in a controlled way by firing the attached HEM package. Alternatively, the partial contribution of the atmospheric drag could be exploited to perform a cost-free altitude lowering. In this respect, the primary propulsion system of the multi-removal space platform could be used to leave the targeted debris into an orbit where the atmosphere effect is significant and a HEM could be attached on it for the atmospheric reentry. In such a way, thanks to the natural decay, the final burn for a controlled reentry will require less Delta-V. This second option introduces the question of in-orbit HEM aging, though a significant mass saving is expected, depending on mission details. With reference to the alternative mission concepts, particular attention will be given to multiple removal feasibility. These aspects will be investigated by using analytical models and numerical simulations. In addition, technological solutions relevant to mission and system aspects will be identified. Results will be used to compare the different mission concepts