Cold Gas Propulsion System Research Articles

Design optimization of cubic-shaped pressurant tank for CubeSat propulsion system

From a technical and economic point of view, cold gas propulsion for nanosatellites is one of the more rational systems. There is a problem, however: the optimal spherical shape for a high-pressure vessel only takes up half the cubic volume of the CubeSat unit. To increase the volume, it is proposed to switch to a cubic-shaped tank. The aim of this work is to synthesize the more rational structural design of a pressurant tank that fits within a cubic constraint, which provides the perfect balance between the mass of the structure and the delta-V budget available for the propulsion system. The proposed variants are as follows: a set of four cylinders with spherical bottoms; 14 equal spheres arranged in a face-centered cubic packing; a thin-walled cube supported by ribs, the location and parameters of which provide the smallest mass of the structure; a cube, the reinforcement scheme of which is calculated using topology optimization method; an original polyspheric and bypass design schemes where the filling of the volume of the cube is created by a set of spherical and cylindrical shells of different radii. All the considered vessel variants can be obtained using additive technologies, such as selective laser sintering. For each variant, the internal volume and mass of the vessels were obtained, as well as the delta-V budget, taking into account the mass of the tank and the mass of the stored gas. The analysis of the results reveals that there are two structures that are more rational. The new bypass scheme provides the largest delta-V budget, but it has a 1.6 times larger mass than the initial spherical vessel design. It should be noted that the difference in vessel mass is not crucial for CubeSat form factor satellites. Choosing a rational design may require a larger delta-V budget. The possibility of increasing the reserve delta-V of the CubeSat cold gas propulsion system by 25–35 % due to the transition from a spherical to a cube-shaped tank has been demonstrated. At the same time, the mass of the tank structure increases by 1.6–6.3 times, which can be considered an acceptable price for increasing the efficiency of the propulsion system. The new bypass scheme offers the optimal balance of mass and delta-V reserve.

SU2 REAL GAS MODELS’ PERFORMANCE PREDICTIONS ON A COLD GAS THRUSTER

Cold gas propulsion systems are preferred, especially for the altitude and trajectory control for satellites, since the 1960s. Both depressurizing the propellant in the propellant tank throughout the mission and the expansion occurring in the divergent part of the thruster nozzle are the reasons for observing very low temperatures and pressure at the outlet section. We have decided to use an open-source compressible CFD tool, SU2, in order to predict the performance of the propulsion system in the early design phase. The scope of this study is the comparison of the results of the vacuum chamber performance tests and the outcomes of unsteady simulations with SU2 using different gas models, including ideal gas, van-der Waals, and Peng-Robinson models. Then, the accuracy and performance of these models are evaluated for the extremely low temperature and pressure conditions. Thruster performance tests have been conducted in the thermal vacuum chamber test campaign at the specially built testing facilities. In order to simulate nominal and low-temperature operation conditions, a propellant tank is thermally conditioned to the predetermined values, and performance test parameters are used as the input for the CFD simulations. The obtained results showed that the van der Waals gas model is the most appropriate gas model for both cases and provides the most realistic results in terms of performance parameters.

Optimization of Micro-Thruster Cold Gas Propulsion System Design for Space Applications

Comparing the complexity of the hardware and the control system, cold gas micro thrusters are much more simplified than other propulsion thrusters. Numerical calculation was carried out for mass requirement, mass flow rate, valves, feed system and computational fluid dynamics with finite element analysis was used for tank, pipe and nozzle design. Four different materials: Structural steel, Stainless steel, Aluminum and Titanium, were considered for tank design. They were subjected to the same conditions for four different tank geometries. The propulsion system was designed to be able to produce between 1 and 2 minutes of continuous thrust and the sudden impulses of between 0.8 and 1 seconds of gas expulsion, with over 20 impulses as desired thrust time. The mass and the volume needed for multiple tank geometries, concepts, and materials were determined. The main performance characteristics of the micro thrusters evaluated were Pressure Distribution, Velocity Profile, Temperature Distribution, Nozzle Efficiency and the corresponding Mach Number Distribution. This work is to provide a stable and controllable platform for testing equipment that can be ultimately applied to space applications.

On a New Type of Combined Solar–Thermal/Cold Gas Propulsion System Used for LEO Satellite’s Attitude Control

This paper presents the development, construction and testing of a new type of solar–thermal propulsion system which can be used for low earth orbit (LEO) satellites. Currently, the vast majority of LEO satellites are fitted with a cold gas propulsion system. Although such a propulsion system is preferred, the service duration of an LEO satellite is limited by the amount of cold gas they carry onboard. In the case of the new type of solar–thermal propulsion system proposed in this paper, the cold gas is first transferred from the main tank in a cylindrical service tank/buffer tank which is placed in the focal line of a concave mirror. After the gas is heated by the solar light focused on the service tank by the concave mirror, it expands by opening the appropriate solenoid valve for the satellite’s attitude control. In this way the service duration of LEO satellite on orbit can increase by 2.5 times compared with a classic cold gas propulsion system. This is due to the propellant’s internal energy increase by the focused solar light. This paper also presents the results achieved by carrying out tests for the hot gas propulsion system in a controlled environment.

Design and development of prototype carpal wrist cold gas propulsion system for attitude control applications

In recent years, the Philippine space program has launched several small-scale satellites, and lately has commissioned its own space program. Research on improving the subsystems of these satellites would be beneficial to future space programs in the Philippines. This paper will detail the process of designing, validating, fabricating, and testing a prototype Carpal Wrist Cold Gas Propulsion System (CW-CGPS). By making use of a carpal wrist’s (CW) hemispherical movement capabilities, the sixteen thrusters could theoretically be replaced by four mounted on opposite ends. Optimization of the propulsion system is based on finding the nozzle design with the most adequate values for thrust, Mach number, and specific impulse while the CW design is validated when it can be calibrated to move as the program dictates. This study uses multiple programs to simulate and verify the design. The conditions of the testing environment are established by resources from international space organizations. Assembly and feasibility are assessed based on the results of the research. It was found that the final optimized system had a model torque of 0.247 N-m, more than enough to overcome the maximum combined influence of the gravity gradient torque of 2.34E-06 N-m and aerodynamic torque of 1.57E-25 N-m. The design, development, and test campaign for the thruster system is presented.

Orbital Maintenance of a Constellation of CubeSats for Internal Gravity Wave Tomography

The advent of radio occultation (RO) instruments aboard CubeSats leads to the possibility of a mission to sound atmospheric internal gravity waves if such satellites are deployed in close-flying constellation. The satellites in the constellation must have slightly perturbed orbital inclinations in order to spread the RO soundings within clusters in two horizontal dimensions, and consequently the satellites will disperse because they will experience different rates of regression of nodes. This dispersion must be countered by propulsive maneuvering in order to maintain the close formation of the constellation. Here, a theoretical approach to the necessary propulsive maneuvering is presented and simulations using comprehensive orbit propagators are performed to analyze four propulsive systems: two cold gas propulsion systems and two electrospray propulsion systems. Cold gas propulsion permits greater separations in inclination between satellites in a constellation by virtue of the greater thrust they can exert on a spacecraft: cold gas propulsion can permit inclination separations of 1 to 10 $^\circ$ while electrospray limits separations to less than 0.2 $^\circ$ . On the other hand, electrospray propulsion provides much longer mission lifetime by virtue of the greater total thrust it offers: cold gas propulsion expends all of its fuel in maintaining the constellation formation in less than approximately 100 days while electrospray propulsion can maintain formation for greater than 1000 days before expending all of its fuel. Mission lifetime is the most critical consideration for a mission, thus electrospray propulsion is recommended for the constellation-flying of CubeSats, but the accelerations that they offer must be greatly increased to enable spacecraft separations useful for tomography of internal gravity waves. Note that any close-flying constellation involving satellites with slightly perturbed inclinations will experience the same dispersing effect as the constellations described herein and, thus, require the same propulsive maneuvering to maintain formation.

Fabrication and Testing of the Cold Gas Propulsion System Flight Unit for the Adelis-SAMSON Nano-Satellites

Adelis-SAMSON is a nano-satellite mission aimed at performing geo-location of target signals on Earth using a tight three-satellite formation in space. To maintain formation, each nano-satellite is equipped with a cold gas propulsion system. The design, qualification, and integration of the Adelis-SAMSON nano-satellite propulsion system is presented in this paper. The cold gas propulsion system mass is approximately 2 kg, takes a volume of 2U, and generates a thrust of 80 mN from four thrusters using krypton as a propellant. We first present the propulsion system requirements and corresponding system configuration conceived to meet the mission requirements. Subsequently, we present the system architecture while listing all the components. We overview the particular role and qualification process of four of the propulsion system’s components: the propellant tank, thruster assembly, pressure regulators, and fill and vent valve. We detail the tests performed on each component, such as proof pressure tests, vibration tests, and external leak tests. Finally, we present the propulsion system level tests before delivery for satellite integration and discuss the propulsion system’s concept of operations.

Development of a Laval nozzle for a cold gas propulsion system

This paper presents a study regarding the calculation and design of a Laval nozzle, and its influence on the performances obtained. For this research, numerical simulation were conducted using the commercial software ANSYS CFX to verify the geometry and to compare it with the analytical results. This study is part of a research project that develops an advanced solar thermal propulsion system; the main purpose is to increase the operation life of a satellite using focused solar light. Two geometries are analyzed in this paper: one with a cone shape and the other generated using the characteristics method. The working fluid is nitrogen gas, which was used due to its inert properties and high molar mass. For the numerical simulations an exterior domain has been taking into account, to capture the jet and to observe how is evolving into the atmosphere. Several cases were studied; temperature was the main parameter varied from one case to another. As a result of the numerical simulations, it was found that the nozzle shape and temperature at the inlet to the convergent section influences performance of the propulsion system in a significant manner. Mach diamonds form because of the supersonic exhaust of the nozzle, and can be seen how their intensity is decreasing in magnitude near the domain outlet.

Cold gas propulsion microthruster for feed gas utilization in micro satellites

Micro spacecraft, which have gained huge popularity in the last decade, are termed as “CubeSats”, a well-known class of small satellites. The fast growing functionality and popularity of CubeSats have helped researchers push technology demonstration towards efficient performance and reliability needed for commercial and governmental applications. Keeping in mind the increasing space debris, dependence of satellite life on fuel, and the use of fossil fuels as propellants, recent efforts are being made to develop a cold gas propellant-based (CGP) micro propulsion system. The CGP system has various merits, namely, non-toxicity, easy to use, and low leakage concerns over other propulsion systems. Liquid propellant-based propulsion requires a liquid feed system, which uses pressurized gas (such as helium, argon and nitrogen) to pressure-feed liquid to the combustion or vaporization chamber so as to produce thrust. But once the liquid propellant is depleted, the pressurized gas is unusable while the system runs out of fuel. Thus, the proposed CGP system being studied, apart from being used as a normal cold gas propulsion system, also offers the advantage of being used as a propulsion system for a liquid fueled system, in the event the liquid fuel runs out. The present work reports on cold gas micro thruster development, utilization of feed gas in a liquid fueled thruster, and the experimental study on the CGP system in simulated vacuum conditions and under a range of pressure conditions. Pressure in the pressure feed gas system for a 1–50 kg category of satellites varies from 4 to 8 bar depending on the mission requirement. The pressure reduces as the liquid propellant gets exhausted. The experimental results are used to validate a computational fluid dynamics model of the system. Thrust values are obtained between the micro to milli-Newton range so as to fulfil the requirements of attitude and station keeping for CubeSats in the 1–50 kg dry mass range. Under vacuum conditions, we obtained thrust values of 0.8 mN at 1 bar feed pressure to 2.24 mN at 4 bar feed pressure. These thrust values are nearly twice those achieved for sea level tests for the corresponding feed pressure. Furthermore, the parameters such as Mach number, velocity vector, pressure, temperature, specific impulse, and nozzle efficiency are studied and reported for atmospheric and vacuum conditions.

Time Optimal State Feedback Control with Application to a Spacecraft with Cold Gas Propulsion

A cold gas propulsion system is well suited to provide the required thrust for a small surveyor spacecraft operated near an asteroid or planetary surface. The cold gas propellant can obtained in-situ from local surface or atmospheric constituents. For small spacecraft, the cold gas system may be limited to only on-off control of the main tank where the generated thrust is directly dependent on the tank pressure. As such the thrust will slowly decrease as the propellant is expended. A state feedback, time optimal, control law is developed for a vehicle with decreasing thrust in translational motion. The success of the control law is shown in simulation.

Modular Testbed for Spinning Spacecraft

The ground verification of a spacecraft control algorithm is commonly done via air bearing facility. Air bearing testbeds are frequently developed for testing a three-axis stabilized spacecraft control algorithm but hardly for a spin-stabilized spacecraft. A modular testbed for testing a spinning spacecraft has been developed at Surrey Space Centre initially for the real-time verification of a prolate spinner slew control algorithm. This testbed is made from commercial off-the-shelf components with a modular system design approach through rapid control prototyping using Matlab xPC Target and is extendable to other rapid control prototyping techniques. It is equipped with a novel low-cost monocular vision system for attitude determination with accuracy of 0.06 deg and angular velocity accuracy of . For the current specification, a cold-gas propulsion system is fixed to the testbed with a two-degree-of-freedom thruster set that can deliver up to 0.25 N of thrust and an air bearing capability that gives three degrees of freedom with a maximum tilt angle of 30 deg. In this paper, the testbed implementation is described, and the test platform is verified.

Designing Space Cold Gas Propulsion System using Three Methods: Genetic Algorithms, Simulated Annealing and Particle Swarm

Cold gas propulsion system technology is relatively less expensive than other propulsion systems. In this paper, for five usual gas as propellant and eight genera as structural used in this systems, first Isp (specific impulse) was checked and then, using three optimization methods, i.e. genetic algorithms, simulated annealing and particle swarm, the optimal total mass of system, principles of structure design, optimal chamber pressure and radius propellant tank were calculated as a multidisciplinary for each gas and structure. Using the results, it is possible to determine the gas, the suitable structure and configuration of the design; for this example, the results of the three algorithms were compared. Assuming that the amount of Thrust, working time and the size of thruster cylindrical chamber is given, it is evident that the gas which have the most Isp is not necessarily a good criterion to be used in the design stage of the cold gas system, and other criteria such as the total mass of the propulsion system, chamber pressure, the radius of the tank and the size of the structures must also be considered.

우주비행체 기동 및 자세제어 설계 검증을 위한 지상 시뮬레이터용 냉가스 추진시스템의 개념설계

최근에 우주비행체의 기동 및 자세제어 로직을 검증하기 위해 소형화 및 정밀화된 부품들을 사용하여 지상에서 시뮬레이터의 구동을 통해 검증하는 연구가 활발하게 진행되는 추세이다. 지상 환경에서 우주비행체 시뮬레이터의 기동 및 자세제어를 위해 일반적으로 구성이 간단하고 신뢰도가 높은 냉가스 추진시스템이 사용된다. 본 연구에서는 우주비행체 지상시뮬레이터의 냉가스 추진시스템 개념설계 결과로서 임무 요구조건에 부합하는 주요 설계 파라미터를 도출하였으며, 전체 시스템 구성을 위한 적절한 사양의 상용부품을 선정하였다. Recently, a validation research of maneuvering and attitude control logics of a spacecraft under a ground condition is getting increase by using operating simulators with compact and precise components. For that, a cold gas propulsion system is generally used for maneuvering and attitude control of spacecraft ground simulators for its simplicity and a high reliability. In the present study, major design parameters of a cold gas propulsion system are derived to meet mission requirements based on conceptual design results of a simulator. And additionally, commercial components with proper specifications are selected for system assembly.

Micro diffuser flow modeling for cold gas propulsion systems

AbstractThe use of highly diluted and hypersonic gas flow is in the scope of application of cold gas thrusters for space applications. Satellites and small spacecrafts are navigated to their orbital trajectory with these nozzles. Inside these propulsion systems high density gradients are dominating the efficiency and the thrust steering behavior of the propulsion systems.Micro flows in the transient regime between free molecular flow and continuous flow are not able to be computed with trustworthy results by using a continuous model with no‐slip boundary conditions. Therefore boundary slip‐velocity models are used for modeling the reduced wall shear stress. Molecular shear stresses decrease the molecular mean velocity near the wall. With a Knudsen number depending slip‐velocity model the effective shear stress is computed by the mean gradient of the velocity profile near the wall.In the present study a trans‐sonic nozzle flow is computed by using a calibrated velocity slip model what depends on the Knudsen number. The Knudsen numbers are lower the Kn=1 at the nozzle neck of the propulsion system. The results are compared with simulation results of a uniform channel flow and computations of the corresponding no‐slip approach. The differences in the hypersonic region are following discussed. (© 2014 Wiley‐VCH Verlag GmbH & Co. KGaA, Weinheim)

Spacecraft attitude dynamics simulator actuated by cold gas propulsion system

This article describes the details of an advanced Spacecraft Attitude Dynamics Simulator (SADS) at the Space Research Laboratory (SRL) at the K. N. Toosi University of Technology. This dumbbell style simulator is based on a spherical air-bearing and employed to develop, improve, and carry out operational tests of sensors, actuators, and attitude control algorithms in experimental framework. The SADS facility includes a variety of components: cold gas propulsion system, complementary inertial measurement unit, on-board processor, semi-automatic mass balancing mechanism, and power supply unit. The overall design of SADS and its components is pretty complicated when considering the mission requirements, operational constraints, and functional limitations imposed by construction procedures. To address this complication, an accurate design and development (or selection) process for each subsystem is presented, which attempts to consider the subsystems interactions and improve cost of operation. The SADS facility is now operated in SRL and the results of its simulated and operational maneuvers are described in detail.

Multi-objective trade-off analysis of an integrated cold gas propulsion system

The overall design of cold gas propulsion systems is pretty complicated when considering the mission requirements, operating constraints and functional limitations imposed by the mechanical components. To address this complication, a precise design process is proposed, which attempts to optimize the cost of operation as well as to minimize the waste volume and weight by using multi-objective trade-off analysis. This analysis is based on a set of ordinary differential equations that are solved iteratively to describe the optimal behavior of the system. Therefore, a numerical code is being developed to give insight on the design sensitivity with respect to uncertainties on the design variables. Consequently, the most favorite operating media, nozzle inlet pressure and tank charging pressure are determined. Moreover, the advantages of electric heater for the storage tank are studied for a specified mission. A three-dimensional computational fluid dynamics analysis is employed to accurately simulate the flow conditions inside the nozzle. The accuracy of the simulation is verified by correlation with experimental results obtained from a non-vacuum test set-up. The experimental results have revealed acceptable performance and agreement with the simulation, within 7% in error.

Simulations of the FORMOSAT-5 Cold Gas Propulsion System by Using the Hybrid Continuum-Particle Method

Reaction control subsystem (RCS) is an onboard satellite propulsion system used to provide required thrusting for orbit raising, orbit maintenance, and attitude control etc.. High pressure flow with high temperature could be generated in a chamber by chemical reactions or other power resources then expelled through a convergent-divergent nozzle to obtain thrust. In order to optimize the thrusting performance, numerical simulation is an efficient method to study the physics and parameters in design phase. In the current study, a hybrid method coupled continuum and particle methods is proposed to simulate flows involving continuum and rarefied regions. The Navier-Stokes (NS) solver named UNIC is developed by Chen and his coworkers. It employs the cell-centered finite-volume method with a hybrid 2D/3D unstructured-grid topology. The proposed particle code named Parallel DSMC Code (PDSC) is a parallelized solver based on the well-known Direct Simulation of Monte Carlo (DSMC) method, which was proposed by Bird in 1976. The physical domain is decomposed into several regimes and each sub-domain executes the serial DSMC code at different processor for speeding up the computing. A practical cold gas thruster with different chamber pressures is simulated by using this hybrid code to study the potential malfunction of the pressure regulator. Then the curve-fitting thrusting equation can be referred to satellite control operations.

부탄을 액화 연료로 사용한 냉가스 추진 시스템에 대한 연구

소형 인공위성의 전형적인 추진 시스템인 냉가스 추진 시스템에 액화 연료를 직접 적용하여 성능을 분석하였다. 고려하는 액화 연료 냉가스 추진 시스템과 일반적으로 사용되는 질소 추진 시스템의 성능을 비교하였다. 질소 추진 시스템과 동일한 질량 조건, 동일한 부피 조건, 동일한 총 임펄스 조건에서 각각 액화 연료를 사용한 냉가스 시스템의 성능과 필요한 연료 탱크의 부피, 필요한 추진 시스템의 질량을 산출하였다. 액화 연료를 사용한 냉가스 추진 시스템은 일반적인 질소 추진제를 사용한 시스템보다 성능, 부피 및 질량 등에서 많은 이점을 가지며 냉가스 추진 시스템에 직접적으로 적용이 가능함을 알 수 있었다. A direct application of liquefied gas propellants to a typical small satellite cold gas propulsion system was analyzed. Performance of systems using liquefied gas propellant under consideration was compared to that of a nitrogen cold gas propulsion system. Liquefied gas propellant propulsion system's performance, required tank volume, and required propulsion system mass has been calculated at the same mass, volume, and total impulse condition of a typical nitrogen cold gas propulsion system. It was found that the liquefied gas propulsion system has advantages in performance, volume, and mass, compared to a typical nitrogen cold gas system, and can be directly applied to a cold gas propulsion system.

Wheel Attitude Cancellation Thruster Torque of LEO Microsatellite during Orbital Maintenance

A cold gas propulsion system is used for orbital maintenance on board microsatellite. Cold gas thrusters are the simplest way of achieving thrust. A microsatellite could be a part of the constellation and to maintain a daily coverage, it will be equipped with a propulsion system for an orbit control. A constellation of several microsatellites could be launched and put at the allocate position in the orbit. To do this, the satellites need few months to be in their final position. A propulsion system is used, among other things, to maintain the satellite at its nominal position. The wheels (reaction/momentum) will be used to dump the thruster disturbances caused by misalignment. This study describes the wheel attitude damping thruster disturbances of Low Earth Orbit (LEO) microsatellite for orbit maintenance with the following points: 1) Attitude dynamics, 2) External disturbances, 3) Magnetic wheel control, 4) Simulation results will be presented to evaluate the performance and design objectives. © 2006 Asian Network for Scientific Information.

Cost effective propulsion systems for small satellites using butane propellant

This paper will describe the work performed at the Surrey Space Centre to produce cost effective propulsion systems for small spacecraft with relatively low deltaV (ΔV) requirements. Traditionally, cold gas nitrogen systems have been used for this type of application, however they have high storage volume requirements. This can be a problem on small spacecraft, which are typically volume limited. An alternative solution is to use liquefied gases, which store as liquids, hence have reasonable density levels, and can be used in a cold gas thruster. At the Surrey Space Centre, butane has been selected as the propellant of choice. Although it has slightly lower specific impulse performance than nitrogen, it has a significantly higher storage density and it stores at a very low pressure, hence no regulation system is required. On 28 th June 2000 Surrey Satellite Technology Ltd (SSTL) launched it first nanosatellite SNAP-1. This 6.5kg spacecraft was equipped with a small cold gas propulsion system utilising 32.6 grams of butane propellant. During the propulsion system operation phase the spacecraft's semi major axis was raised by nearly 4 kilometers using the propulsion system. The design of the propulsion system will be described and the low cost features highlighted. Telemetry data will be used to describe the propulsion operations and an overall mission specific impulse will be derived. SSTL are currently under contract to build three Earth observation spacecraft for a Disaster Monitoring Constellation (DMC). Each spacecraft will weigh approx 100 kg and have a ΔV requirement of 10 m/sec. A butane system has been designed and manufactured to meet the requirements of these spacecraft. The system is based very much on the flight heritage of the SNAP-1 system, with the addition of greater propellant storage capacity. The lessons learnt from the SNAP-1 operation will be reviewed and the resulting design improvements on the DMC propulsion systems will be detailed.