Cooling Rocket Motor Apparatus

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System, method and apparatus for cooling rocket motor components using a saturated liquid vapor coolant mixture Abstract A system and method of cooling a rocket motor component includes injecting a high pressure liquid coolant through an injector nozzle into a cooling chamber. The cooling chamber having a pressure lower than the high pressure liquid coolant. The liquid coolant flashes into a saturated liquid-vapor coolant mixture in the cooling chamber. The saturated liquid-vapor coolant mixture is at equilibrium at the lower pressure of the cooling chamber. Heat from the rocket motor component to be cooled is absorbed by the coolant. A portion of the liquid portion of the saturated liquid-vapor coolant mixture is converted into gas phase, the converted portion being less than 100% of the coolant. A portion of the coolant is released from the cooling chamber and the coolant in the cooling chamber is dynamically maintained at less than 100% gas phase of the coolant as the thrust and heat generated by the rocket motor varies. Description The present invention relates generally to rocket motors, and more particularly, to methods and systems for cooling rocket motors and components of rocket motors. Liquid fuel rocket motors are traditionally cooled using the liquid fuel. The liquid fuel is circulated through a cooling chamber around portions of the rocket motor the need to be cooled such as the combustion chamber and the outlet. Liquid fuel rocket motor uses a liquid fuel and a liquid oxidizer that when combined in the combustion chamber produces thrust and of course tremendous amounts of heat. The liquid cooling is provided to extend the service life of the rocket motors. Without liquid cooling the rocket motor would typically erode or burn away the inner surfaces of the rocket motor. Thus rendering the rocket motor as a single use rocket motor or inoperative or even resulting in a catastrophic failure such as an explosion. Some of the more modern rocket motors are hybrid rocket motors. Hybrid rocket motors use a solid fuel and a gas or liquid oxidizer. They are termed hybrid rocket motors because the fuel and the oxidizer are in different material phases i.e. solid phase and liquid or gas phase. Hybrid rocket motors are typically simpler than a liquid fuel rocket motor while also providing much of the same operational advantages of a liquid fueled rocket motor (e.g., throttling, multiple firing and shutdown cycles, etc.). In a hybrid rocket motor the solid fuel cannot be circulated through the cooling chamber around portions of the rocket motor because the fuel is solid phase. Using a liquid oxidizer as the rocket motor coolant has substantial limitations because the heat absorbed into the oxidizer from the rocket motor would require the oxidizer to be maintained at a very high pressure in the cooling chamber. Further, most liquid oxidizer's are not very efficient coolants. Further still, a heated oxidizer can be very difficult to manage safely. As a result rocket the motor would be more likely to have a catastrophic failure. Thus reducing the overall reliability of the hybrid rocket motor which is the precisely opposite goal of the hybrid rocket motor. However there is still a need for cooling hybrid rocket motors to extend the life of the hybrid rocket motor so that they can be used multiple cycles and multiple flights. Summary Broadly speaking, the present invention fills these needs by providing a system, method and apparatus for cooling hybrid rocket motors. It should be appreciated that the present invention can be implemented in numerous ways, including as a process, an apparatus, a system, computer readable media, or a device. Several inventive embodiments of the present invention are described below. A system and method of cooling a rocket motor component includes injecting a high pressure liquid coolant through an injector nozzle into a cooling chamber. The cooling chamber having a pressure lower than the high pressure liquid coolant. The liquid coolant flashes into a saturated liquid-vapor coolant mixture in the cooling chamber. The saturated liquid-vapor coolant mixture is at equilibrium at the lower pressure of the cooling chamber. Heat from the rocket motor component to be cooled is absorbed by the coolant. A portion of the liquid portion of the saturated liquid-vapor coolant mixture is converted into gas phase, the converted portion being less than 100% of the coolant. A portion of the coolant is released from the cooling chamber and the coolant in the cooling chamber is dynamically maintained at less than 100% gas phase of the coolant as the thrust and heat generated by the rocket motor varies. Other aspects and advantages of the invention will become apparent from the following detailed description, taken in conjunction with the accompanying drawings, illustrating by way of example the principles of the invention. Detailed Description Several exemplary embodiments for a system, method and apparatus for cooling rocket motors will now be described. It will be apparent to those skilled in the art that the present invention may be practiced without some or all of the specific details set forth herein. One approach to cooling rocket motors and components thereof is to use the phase change of a liquid coolant to cool the rocket motor. One or both of the fuel and the oxidizer could be used as the coolant in a liquid fueled rocket motor. If both the oxidizer and the fuel are used to cool the rocket motor, each of the oxidizer and the fuel are contained in separate cooling chambers that are thermally coupled to the portions of the rocket motor to be cooled. In a hybrid rocket motor a liquid oxidizer could be used as the coolant. The phase change from liquid to gas requires a quantity of energy known as the heat of vaporization. The heat of vaporization is different for each coolant material. As the coolant absorbs the energy to meet the heat of vaporization and change the state of the coolant from liquid phase to gas phase, the temperature of the coolant does not increase. Thus, the coolant can be maintained at a nearly constant temperature while absorbing energy in the form of heat from the rocket motor. The heat of vaporization is the amount of heat required to convert a given mass of a material in its liquid phase into the gas phase at constant temperature and pressure. The gas phase material will release the same amount of heat when it condenses to become liquid phase. In one approach, the coolant can be injected into the cooling chamber in a saturated liquid-vapor form. A saturated liquid-vapor form is when the liquid phase coolant and gas phase coolant are in equilibrium for the present temperature and pressure. The heat from the rocket motor converts the liquid phase portion of the saturated liquid-vapor coolant to the gas phase of the coolant. In this way the temperature of the coolant in the cooling chamber does not substantially change. Instead the heat energy absorbed from the rocket motor is used to change the phase of the coolant from liquid phase to gas phase. The amount of the heat the saturated liquid-vapor coolant mixture can absorb from the rocket motor is a function of a residence time of the coolant in the cooling chamber, a volume of the coolant in the cooling chamber and the heat of vaporization of the coolant. The gas phase coolant that is output from the cooling chamber can be vented off or used in the combustion process. By way of example, a liquid phase of the coolant at a first pressure is injected into the cooling chamber through an injector. The pressure of the cooling chamber is a second pressure lower than the first pressure. Therefore, the liquid coolant will drop to the second pressure and flash to form a saturated liquid-vapor coolant mixture at an equilibrium point for that temperature and pressure. The heat from the rocket motor is absorbed by the liquid phase portion of the saturated liquid-vapor coolant mixture in the cooling chamber. This heat satisfies the heat of vaporization of at least a portion of the liquid phase portion of the saturated liquid-vapor coolant mixture to convert to a gas phase. The gas phase coolant is output from the cooling chamber before the coolant in the cooling chamber is 100% gas. Any coolant having a suitable heat of vaporization could be used to cool the rocket motor. One example of a suitable coolant is nitrous oxide (N2O) which is also used as an oxidizer. It should be understood that other oxidizers and fuels could also be used as rocket motor coolants. Using an oxidizer as the cooling media in the hybrid rocket motor provides several benefits. First, the oxidizer is already onboard the rocket for oxidizing the fuel and using the oxidizer cooling eliminates a requirement of carrying a third material for coolant purposes. Second, the oxidizer used as the coolant can then be consumed in the combustion to oxidize the fuel. In a more specific example of the rocket motor cooling process, a saturated liquid-vapor coolant mixture is injected into the cooling chamber. At 100% gas phase of the coolant, the coolant cannot absorb any more heat from the rocket motor without increasing in at least one of temperature or pressure. Further, the gas phase of the coolant is not nearly as efficient absorbing rocket motor heat as the liquid phase of the coolant. In one embodiment the temperature and pressure of the cooling chamber can be monitored and the coolant can be output or flow through the cooling chamber increased if there is a significant pressure and/or temperature increase in the cooling chamber as that would indicate the coolant in the cooling chamber has reached or is approaching a condition of 100% gas phase of the coolant. A rapid rate of increase in pressure or temperature is another mechanism for indicating the coolant is approaching a condition of 100% gas phase of the coolant. The rapid rate of increase in temperature or pressure would be greater than of a rate of increase in temperature of the rocket motor component being cooled by the cooling chamber. By way of example; the rocket motor component being cooled might experience a low rate of increase in temperature of 10 degrees over a 10 second period if the temperature of the coolant in the cooling chamber increased more than 10 degrees in 10 seconds, then the coolant could be approaching 100% gas phase. As a result, a coolant flowrate could be increased. If additional heat were added to the 100% gas phase of the coolant, the coolant can begin to dissociate and break down into its constituent elements. By way of example, a nitrous oxide coolant would break down into elemental nitrogen and oxygen. Dissociating nitrous oxide is an exothermic reaction and therefore releases additional heat. This additional heat can cause a cascade of dissociation and exothermic reaction resulting in a potentially catastrophic overheating of the rocket motor as well as a rapid increase in pressure in the cooling chamber and a potential catastrophic mechanical failure of the cooling chamber and the rocket motor due to over-pressurization. Therefore, the residence time of the saturated liquid-vapor coolant mixture in the cooling chamber is limited so that the coolant in the cooling chamber does not achieve 100% gas phase of the saturated liquid-vapor coolant mixture. As stated above, the coolant output from the cooling chamber can be injected into the combustion chamber for combustion with the fuel and/or vented. One use of the cooling system and method described here in is in cooling rocket motor components. One of the difficulties with hot-fired aerospike rocket motor nozzles, is the high heat loads at the base of the spike, which can cause ablation of the spike and a degradation of performance, and eventually even cause failure of the entire nozzle and rocket motor. The purpose of a rocket motor nozzle is to accelerate a propellant, from a reservoir at high stagnation pressure, for the purpose of creating thrust. The effectiveness with which the rocket motor produces thrust is generally characterized by the nozzle thrust coefficient. The nozzle thrust coefficient is thrust normalized by the product of throat area and nozzle stagnation pressure. Definitions of symbols used herein: Ae=Nozzle exit plane area A*=Nozzle throat area CF=Thrust Coefficient Dt=Throat diameter h=Local heat transfer coefficient Isp=Specific impulse Pe=Exit plane pressure P0=Nozzle stagnation pressure P∞=Ambient pressure R=Axial radius of curvature of nozzle Tcu=Measured copper throat temperature (cold side) z=Axial distance along nozzle γ=Ratio of specific heats Equation (1) expresses the thrust coefficient as a function of the rocket motor operating parameters and of the ambient pressure into which the nozzle discharges. C F = 2 ⁢ γ 2 γ - 1 ⁢ ( 2 γ + 1 ) ( γ + 1 ) / ( γ - 1 ) ⁡ [ 1 - ( P e P 0 ) ( γ - 1 ) / γ ] + ( P e - P ∞ P 0 ) ⁢ A e A *                            Equation ⁢ ⁢ 1 The nozzle performance, as characterized by the thrust coefficient, is maximized when the nozzle operates at fully expanded conditions, i.e., when the nozzle exit plane pressure equals the ambient pressure, a condition that defines the nozzle's design pressure ratio P0/P∞. For a conventional converging-diverging nozzle, with fixed area ratio, this means that performance can only be optimum at a single operating ambient pressure. For the continuously-changing ambient pressure (and thus nozzle pressure ratio) of a typical rocket atmospheric trajectory, the thrust coefficient is therefore non-optimum for much of its operation.