Browse Topic: Propellants
To solve a problem that ignition anomaly can’t be detected in time, based on the thermal equilibrium equation, the space heat flow, heater heating, propellant combustion, and thermal radiation to cryogenic space are considered to build an accurate ignition temperature method for the 10 N thruster by using on-orbit true temperature. Further, considering the error of measuring the thermistor, an envelope model for the 10 N thruster ignition temperature is established. Based on the above, a detection method for the 10 N thruster ignition anomaly of on-orbit satellites is proposed. The accuracy of the method is relatively high, and the absolute error is less than 3 degrees Celsius. An anomaly can be quickly detected when the 10N thruster ignition temperature deviates from the normal trend by 3–5 degrees celsius. The method is applied to a DFH-3 satellite, and the maximum difference of 10 N thruster ignition temperature between the theoretical values calculated by the proposed method and the measured values is only 2.72 degrees celsius. It has been proven that the prediction accuracy of the proposed method is high. It plays an important role in discovering the 10N thruster ignition anomaly in time and ensuring the success of satellite orbit or attitude control.
Sealing systems in space applications must perform reliably under demanding conditions in engineering: cryogenic temperatures, vibration, leakage control, ultra-high vacuum, ionizing radiation, abrasive particulates, and repeated thermal cycling. Each factor strains conventional sealing technologies. In combination, they can rapidly cause failure in systems where margins are unforgiving and maintenance is impossible. As spacecraft architectures evolve toward longer operational lifetimes and broader mission profiles, sealing requirements continue to tighten. Launch vehicles, satellites, and exploration platforms now operate across wider temperature ranges and in contact with more aggressive propellants and media. As a result, both metal seals and engineered polymer alternatives are evaluated-and selected-against increasingly specific, measurable performance criteria.
Plastic materials are used for a wide variety of spacecraft applications including seals, bearings, fasteners, electrical insulators, thermal isolators, and radomes. Selecting plastics for use in space is complex due to wide operating temperature ranges, vacuum conditions, and exposure to radiation and atomic oxygen. Additionally, some spacecraft applications require sealing flammable propellants such as hydrogen and oxygen. This article will present some design considerations when selecting plastics for use in spacecraft. It will provide rich data on the performance characteristics of plastics as well as examples of successful spacecraft applications.
The race is on for leadership in cislunar space, considered a gateway to the future of space exploration. Yet operating in this domain introduces unique challenges for propulsion systems. In contrast to low-Earth orbit (LEO), the cislunar environment requires higher precision propulsion solutions; these are necessary to enable rapid and accurate maneuvering of spacecraft and long-term sustainability. Propellants like hydrazine and nitrogen tetroxide offer the high energy density required for cislunar missions, but they must be handled very differently from the inert, non-reactive gases at play in LEO systems.
Engineers at NASAs Stennis Space Center have developed the HYdrocarbon Propellants Enabling Reproduction of Flows in Rocket Engines (HYPERFIRE), a sub-scale, non-reacting flow test system. HYPERFIRE uses heated ethane to enable physical simulation of rocket engines powered by a broad range of propellants in an inexpensive, accurate, and simple fashion.
As NASA plans for manned missions to Mars, efforts are being made to identify technologies that must be improved to make such trips feasible. Without improvements in valve technologies, propellant and commodity losses will likely make long-duration space missions infeasible. Engineers at NASA's Marshall Space Flight Center have developed a self-aligning poppet for low leakage valves a seat alignment technology that eliminates the need to precisely control interfaces between poppet sealing surfaces and the valve seat seal.
NASA Goddard developed the Cooperative Service Valve (CSV) to facilitate the resupply of media, such as propellants and pressurants, to satellites. The CSV replaces a standard spacecraft fill and drain valve.
Researchers have used the ancient Japanese art of paper folding to possibly solve a key challenge for outer space travel: how to store and move fuel to rocket engines. They developed an origami-inspired, folded plastic fuel bladder that doesn’t crack at super-cold temperatures and could someday be used to store and pump fuel.
Strategies to control solid rocket propellant regression rate require a robust throttling technique applicable to high performance propellant formulations. Currently, several methods to control and throttle either motors or subscale propellant strands exist, including chamber pressure control (e.g. pintle nozzles or rapid depressurization quench), infrared laser irradiation of the burning surface to increase burning rates, development of inherently unstable combustion chamber geometries (producing either local pressure or velocity perturbations), and electrically sensitive hydroxylammonium nitrate (HAN)-based formulations in which burning rate is controlled by a voltage potential. However, these techniques are limited in that they either can only be used with low flame temperature (low specific impulse) propellants, result in low propulsion system mass fraction (pintle), are only capable of producing a single perturbation, or are formulation specific.
Researchers have developed a new propulsion concept for swimming robots that exploits temperature fluctuations in the water for propulsion without the need for an engine, propellant, or power supply.
A highly miniaturized, MR-143, green monopropellant thruster was developed for 1N thrust. Testing indicated the initial catalyst bed heater was insufficient. In subsequent development, the thruster was equipped with a more efficient catalyst bed heater. For reliable ignition of the advanced, non-toxic, AF-M315E monopropellant, the catalyst needs to be preheated. This preheat temperature is much higher than what hydrazine thrusters require. Moreover, the combustion temperature of hydroxyl ammonium nitrate (HAN)-based monopropellants is higher than hydrazine, so the catalyst bed heater must be able to withstand repeated soak-back temperatures.
The scope of this SAE Information Report is to provide general information relative to the nature and use of infrared techniques for nondestructive testing. The document is not intended to provide detailed technical information, but will serve as an introduction to the theory and capabilities of infrared testing and as a guide to more extensive references.
Putting a satellite into low Earth orbit requires a lot of energy, with ground-launched rockets expending two-thirds of their propellant fighting to get through the atmosphere. Researchers at NASA’s Armstrong Flight Research Center have developed an innovative approach to launching satellites into space from an airborne platform. As with other air-launch approaches, it provides significant flexibility in the location and direction of the launch vehicle. Furthermore, unlike other air-based launch techniques, this system avoids the significant drawbacks related to expensive and complex design/development efforts, difficult maneuvering, risks to crew, and inefficient flight performance.
Structural analysis of solid rocket motors is challenging for several reasons, but the most important of these is the complex behavior of the propellant. The mechanical response of a solid propellant is time and temperature dependent. The complexity of the mathematical analysis of the propellant depends on the loading conditions, but for some loading situations, the linear viscoelasticity assumption is reasonable. In particular, linear viscoelasticity is perhaps the most appropriate material behavior description for use in the simulations of stresses related to storage conditions. Typically, simulations use a viscoelastic model in the form of a Prony series and a Williams–Landel–Ferry (WLF) equation. The parameters in these models are derived from stress relaxation experiments, making the stress relaxation experiment a key viscoelastic test, analogous to the tensile test for linear elastic materials.
In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously.
Test methodology allows analysis of combustion dynamics for subscale rocket injectors under super critical conditions. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California In the study of combustion characteristics of liquid rocket fuels, it is customary to either study the combustion of liquid fuel droplets or the combustion of fuel sprays. However, the two are closely related to each other, because in a typical rocket combustion chamber, the burning of droplets, droplet clusters, and fuel sprays occur simultaneously. The study of droplet burning is mostly focused either on single suspended droplets or free droplets, usually in a free fall. In either case, the conditions regarding the combustion of a single fuel droplet deviates from what is expected in typical liquid rocket engines (LREs). This not only concerns the chamber pressure and the temperature that is usually much lower in a typical droplet study, but also the droplet ignition process and the related transient effects such as temperature ramp up that is much slower than what takes place in LREs. As a result, reported values of combustion parameters such as the burning rate constant and the ignition delay time obtained under typical experimental conditions may not entirely represent the combustion processes in LREs.
Humans have been using rocket propulsion for almost a millennium, starting with Chinese rockets and “fire arrows” in the 13th century and continuing to the modern era's powerful Space Shuttle and Falcon rockets. For most of that history, rockets have been chemically fueled, but in the past century scientists and engineers have also experimented with electric rockets, also known as ion engines or ion propulsion systems. Rather than using chemical reactions to create heat and accelerate a propellant, electric rockets use electromagnetic or electrostatic fields acting on charged ions of propellant, speeding them up and shooting them out, away from the vehicle, producing thrust. The electrical energy to generate these fields comes from the sun, from batteries, or both.
New testing technique requires less material, gives more accurate results. Air Force Research Laboratory (AFMC), Edwards Air Force Base, California Structural analysis of solid rocket motors is challenging for several reasons, but the most important of these is the complex behavior of the propellant. The mechanical response of a solid propellant is time and temperature dependent. The complexity of the mathematical analysis of the propellant depends on the loading conditions, but for some loading situations, the linear viscoelasticity assumption is reasonable. In particular, linear viscoelasticity is perhaps the most appropriate material behavior description for use in the simulations of stresses related to storage conditions. Typically, simulations use a viscoelastic model in the form of a Prony series and a Williams-Landel-Ferry (WLF) equation. The parameters in these models are derived from stress relaxation experiments, making the stress relaxation experiment a key viscoelastic test, analogous to the tensile test for linear elastic materials. A typical set of stress relaxation tests is performed at several discrete temperatures that cover a range of temperatures anticipated by the fielded motor. At each of the selected temperatures, the specimen is deformed with approximately a single step in strain, which is then held constant for the duration of the test. While held at this constant strain, the stress decays over time due to relaxation of the rubbery elastomer. During this portion of the test, the stresses are measured, and the ratio of stress to applied strain is determined. This ratio is termed the stress relaxation modulus ER. Using time-temperature superposition, the set of curves at the various temperatures can be shifted horizontally relative to each other to form a master curve. The translation of the curves takes a specific mathematical form, viz., the WLF equation. From this master curve, the Prony series at any given temperature can be calculated, and the calculation can be incorporated into finite element analyses along with the WLF equation, making linear viscoelastic analysis of rocket motors possible.
Innovators at NASA's Glenn Research Center have developed several new technological innovations to improve the capability of Hall-effect thrusters, which are used primarily on Earth-orbiting satellites and can also be used for deep-space robotic vehicles. Hall thrusters are susceptible to discharge channel erosion from high-energy ion impingement, which can reduce operational thruster lifetimes. Glenn researchers have developed several approaches to mitigate this problem. One is a magnetic circuit design that minimizes discharge chamber ion impingement. Another successful improvement developed by Glenn is a means of replacing eroded discharge channel material via a channel wall replacement mechanism. A third innovation is a propellant distributor that provides both a high degree of flow uniformity, and shielding from back-sputtered contamination and other potential contaminants. All of these advances work toward increasing the operational lifetime and efficiency of Hall thrusters.
JPL's Microfluidic Electrospray Propulsion (MEP) thruster design is based on a microfabricated electrospray system with a capillary-force-driven feed system that uses indium metal as the propellant. This architecture provides an extremely compact, modular system scalable to a wide range of applications from micro spacecraft to large, space-based telescopes.
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