Propulsion
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O/ F Ratio or Mixture Ratio
Mixture Ratio is the ratio of the liquid oxidizer flow rate divided by the liquid fuel flow rate, with both flow rates being measured as mass flow rates. The best performance (highest specific impulse) is obtained at a specific optimum mixture ratio.
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Regenerative Cooling
Regenerative cooling, in the context of rocket engine design, is a configuration in which some or all of the propellant is passed through tubes, channels, or in a jacket around the combustion chamber or nozzle to cool the engine. This serves two purposes: 1) it cools the combustion chamber and 2) it warms the fuel -- making the combustion reaction more efficient. [1]
Ablative Cooling
With ablative cooling, combustion gas-side wall material is sacrificed by melting, vaporization and chemical changes to dissipate heat. As a result, relatively cool gases flow over the wall surface, thus lowering the boundary-layer temperature and assisting the cooling process.
Film Cooling
Water cooling
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| Word | Definition | Assosciated Links and Resources |
| ------------------------------- | --------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------------- | ---------------------------------------------------------------------------------------- |
| O/ F Ratio or Mixture Ratio | Mixture Ratio is the ratio of the liquid oxidizer flow rate divided by the liquid<br>fuel flow rate, with both flow rates being measured as mass flow rates. The best performance (highest specific impulse) is obtained at a<br>specific optimum mixture ratio. | |
| Chamber Pressure | The Chamber Pressure is the pressure in the combustion chamber of an<br>operating rocket propulsion system. | |
| Propellant | Propellant is the stored matter that is energized and ejected. It is commonly in the form of a Liquid<br>Propellant (stored in vehicle or missile tanks) or a Solid Propellant (stored<br>inside its combustion chamber). Rocket propellants undergo a chemical reaction<br>or a Combustion Process (usually at high chamber pressure and high<br>temperature) that transforms them into hot gaseous reaction products or Exhaust<br>Gases, which are then accelerated and ejected through a Supersonic Nozzle. | |
| Specific Impulse | Specific Impulse is a parameter indicating propulsion system performance. It can<br>be defined as the thrust of an equivalent rocket propulsion system (same chamber<br>pressure, same propellant, same nozzle throat to exit area ratio) that has a<br>propellant mass flow of unity. Higher values indicate a better system. | |
| Total Impulse | Total Impulse is the integral of thrust over the propulsion operating time. It is a<br>measure of the total kinetic energy of the nozzle exhaust gas as released by the<br>combustion of all the available propellant in the propulsion system. For constant<br>thrust operation it is the average thrust multiplied by the effective propulsive<br>operating duration or it is also the mass of the total expelled propellant multiplied<br>by the average specific impulse. | |
| RPA | Rocket Propulsion Analysis: Software tool that helps model out conceptual rockets | [http://www.rocket-propulsion.com/index.htm](http://www.rocket-propulsion.com/index.htm) |
| Cryogenic Propellants | Cryogenic propellants are subcooled liquids at low temperature (such as liquid<br>oxygen or liquid hydrogen); they are gases at ambient temperatures. | |
| Nozzle Area ratio | The Nozzle Area Ratio is the nozzle exit area divided by the nozzle throat area.<br>For optimum gas expansion in a nozzle the gas pressure at the nozzle exit is equal<br>to the local ambient atmosphere pressure. Typical values of this nozzle area ratio<br>are between 4 and 20 for expansion to sea-level pressure and between 40 and 200<br>for operation at very high altitude (space vacuum). | |
| Ablative Chamber or Nozzle Well | An Ablative Chamber or Nozzle Well absorbs heat (from the hot gases) by<br>having some of the heated ablative well material (e.g., reinforced organic fibers in<br>an organic matrix or phenolic resins) vaporized, endothermically decomposed into<br>gaseous species, or charred. The evaporated gas products flow out of the ablative<br>material and form a protective cooler gas layer at the wall’s surface. | |
| Liquid Engine (LE) | A Liquid Propellant Rocket Engine commonly has these principal components: one or two<br>propellant tanks, one or more thrust chambers, a feed mechanism (pumps —<br>driven by a turbine or displacing the propellant in the tank(s) by high pressure<br>inert gas), piping and control valves, and sometimes servo- valves (for starting,<br>stopping, throttling, or mixture ratio control). | |
| Rocket Propulsion Systems | A Rocket Engine (it uses liquid propellants) and a Rocket Motor (solid propellants) are the two most common types of Rocket Propulsion System. | |
| Components of a rocket engine | A Rocket Engine usually consists of one or more Thrust Chambers, one or<br>more Tanks for storing propellants, a Feed Mechanism to force the liquids into<br>the thrust chamber, a Power Source to provide energy to the feed mechanism,<br>suitable Piping and Valves to transfer the liquid propellants, a structure to<br>transmit the thrust force to the vehicle, and Controls to initiate and regulate the<br>propellant flow rates. | |
| Components of a thrust chamber | The Thrust Chamber, also often called Thruster (usually for low thrust and<br>repeated starts), has an Injector (where propellants are introduced and metered), a<br>Combustion Chamber (where liquid propellants react and burn) and a<br>converging-diverging Nozzle, where the hot combustion gas is accelerated and<br>ejected at supersonic velocities. | |
| Solid rocket propellant | Solid Rocket Propellant typically consists of an oxidizer (usually a crystalline<br>solid like ammonium perchlorate), an organic Fuel (such as a rubbery polymer<br>like polybutadiene, which also acts as the glue to hold the grain together), and<br>various additives to improve performance, storage, thrust-time profile,<br>manufacture, aging, etc. Additives include liquid Plasticizers, Explosives,<br>Burning Rate Catalysts, etc. | |
| burn rate | The Burning Rate is the rate of regression of the burning grain surfaces as<br>propellant is consumed or burnt (inches per second) in a direction normal to the<br>surface. Surfaces that are bonded to the case walls or to insulators, will not burn. | |
| inhibitor | Inhibitors are layers of non-burning materials that are glued to exposed grain<br>surfaces so that they will not burn. The propellant flow and, therefore, also the<br>thrust are proportional to this burning rate and the exposed burning surface. The<br>burning rate varies with chamber pressure and the initial ambient temperature of<br>the grain. | |
| internal insulator | Internal Insulators are layers on the inside of the case wall made of material<br>with low thermal conductivity; they protect the case from the hot combustion<br>gases and prevent it from reaching the temperature where the case material loses<br>its strength | |
| external insulator | External Insulators are applied to the outside of liquid propellant tanks or solid<br>propellant motor cases to protect against excessive heat transfer from hot air,<br>when flying through the atmosphere at high speed. | |
| Attitude control | Attitude Control (vehicle rotation) and a limited amount of Flight Velocity<br>Change can also be achieved by a series of small thrusters, which are located in<br>various locations on the vehicle. By Pulsing individual thrusters or pairs of<br>thrusters (repeated short duration operations) the vehicle can be rotated and<br>maneuvered. | |
| Erosion | In uncooled nozzles the gradual Erosion of the nozzle throat causes a small increase in nozzle cross-sectional area and thus a small decrease in chamber pressure and thrust.<br><br> |Last updated