Patent Application
20100326043
1. Field of the Invention
The present invention relates generally to power plants, and more specifically to an expander cycle rocket engine nozzle.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
Rocket engines that use cryogenic fuels and oxidizers such as a gas generator cycle and the staged combustion cycle rocket engines use some of the fuel and oxidizer to pre burn and drive the turbo-pumps that deliver the high pressures to the engine nozzle. In the gas generator cycle, a very small portion of the fuel and the oxidizer is bled off from that delivered to the main combustion chamber (MCC) and diverted into a small pre-burner and combusted to produce a hot gas flow that is then used to drive the turbo-pumps that supply the high pressure fuel and oxidizer to the main combustion chamber. The exhaust gas from the turbo-pumps is then vented overboard. In the staged combustion cycle, a small portion of the propellant (either the oxidizer or the fuel) is diverted and combined with all of the other propellant to partially combust the combination, which is then passed through the turbo-pumps to drive these. This mixture is then sent to the main combustion chamber along with the remainder of the propellant and is combusted in the main combustion chamber. In both the gas generator and staged combustion cycles, some of the fuel and oxidizer is used to produce power to drive the turbo-pumps and therefore not used to produce power in the rocker engine nozzle. Also, because of the high turbo-pump inlet temperature, the turbine driving the turbo-pump is subject to thermal shock and thermal mechanical failure, or TMF.
An expander cycle rocket engine passes a propellant (typically the fuel) through a heat exchanger formed within or around the nozzle to transfer heat from the combustion process in the nozzle to the fuel to heat up the fuel. The heated fuel (in the case of most cryogenic fuels and oxidizers is hydrogen) is passed through the turbines that drive the turbo-pumps to pressurize the fuel and the oxidizer prior to injection into the main combustion chamber for combustion. The expander cycle is more efficient than either of the gas generator and staged combustion cycles because all of the fuel and oxidizer is used in combustion and exhausted through the throat and then into the nozzle for expansion. The expander cycle rocket engine has many advantages over the staged combustion and gas generator cycles. Rather than using a pre-burner, the engine routes liquid propellant from the pump discharge to the nozzle. This flow cools the nozzle and heats up the liquid turning it into a gas. The high pressure gas is then routed to the turbine inlet to drive the turbo-pump(s). The turbine is driven by gas expanded from heat transfer in the engine nozzle rather than from products of combustion from a pre-burner as used in the gas generator and staged combustion engine cycles. As a result, the turbine temperature is significantly lower than for the other cycles resulting is longer life due to the elimination of thermal shock and thermal mechanical fatigue (TMF). The expander cycle rocket engine has proven to be the most reliable engine and has demonstrated superior re-start capability. However, in prior art expander cycle rocket engines, the thrust this engine is capable of producing has reached a maximum limit. As the size of the nozzle increases, the engine mass flow increases at a greater rate than the surface area of the nozzle. As a result, a limit is reached when there is insufficient heat transfer in the nozzle to drive the turbo-pump(s) required to provide the mass flow to the engine. Additionally, for a given mass flow, the chamber pressure is also limited based on the turbine power available for driving the turbo-pumps
High thrust (in excess of 100,000 pounds) expander cycle rocket engines have traditionally been limited to a chamber pressure below 1,500 psia because of a lack of turbine power available to the fuel turbo-pump. In a typical expander cycle rocket engine, fuel from the fuel turbo-pump is pumped through the cooling liner and tubular nozzle of the engine's nozzle assembly where the fuel is heated and then fed to a turbine which drives the turbo-pumps. In order to increase the combustion chamber pressure, flow to the combustion chamber must be increased. However, as fuel flow through the cooling liner and tubular nozzle increases, the temperature of the fuel at the turbo-pump turbine inlet decreases due to the increase in mass flow rate of the fuel or to provide higher discharge pressure. At the same time, the fuel turbo-pump must do more work to provide the increased mass flow rate of the fuel. Although the energy available to the fuel turbo-pump turbine is a function of both the mass flow rate of the fuel and the turbine inlet temperature, the increase in the mass flow rate of fuel cannot offset the resulting decrease in turbine inlet temperature which occurs as a result of the increased fuel flow rate. Consequently, the decrease in turbine inlet temperature and the increase in work required by the turbo-pump at the higher fuel flow rates act to limit the maximum fuel flow rate to the combustion chamber, thereby limiting combustion chamber pressure.
In summary, the expander cycle rocket engine uses heat from the nozzle to heat up the fuel to drive the turbo-pumps that pressurized the fuel and oxidizer for combustion in the nozzle (combustion chamber). To increase the thrust of the rocket engine, a larger propellant flow and/or discharge pressure is required. As the engine/nozzle size increases, the propellant volume increases faster than the surface area of the nozzle. As the nozzle volume increases and more fuel and oxidizer is needed to be pressurized, the amount of heat transferred to the fuel for driving the turbo-pumps becomes less than required to supply the higher pressures. As a result of increasing the nozzle volume, the efficiency of the expander cycle rocket engine becomes less and less. There is a limit in nozzle size using the present technologies because of this effect.
It is therefore an object of the present invention to provide for an expander cycle rocket engine that can have a higher thrust or a higher chamber pressure than available in the cited prior art.
The present invention is nozzle assembly for an expander cycle rocket engine in which the nozzle assembly includes an primary nozzle of prior art design and an inner nozzle shaped heat exchanger formed concentric with the primary nozzle, and in which both the primary nozzle and the nozzle shaped heat exchanger include heat transfer passages in which the fuel is passed to cool the heat exchanger and to extract heat from the combustion gases in the nozzle. Because of the concentric nozzle arrangement, more heat can be extracted from the combustion gases and used to drive the turbo-pumps to supply the required higher pressure to the fuel and oxidizer supplied to the main combustion chamber. The inner nozzle shaped heat exchanger is secured in place by flow entrance and exit struts that pass the fuel between the primary nozzle and the nozzle shaped heat exchanger (the inner nozzle). The inner nozzle conforms to the flow of the combustion gas within the primary (outer) nozzle so as not to block the exhaust gas flow.
The present invention is a nozzle assembly for use in an expander cycle rocket engine.
In an alternate embodiment, cross fins 15 can be includes within the inner nozzle 13 that also function to pass the liquid fuel to extract more heat from the combustion gases flowing through the nozzles. The cross fins 15 can be used with the inner nozzle 13 and extend between the inner nozzle 13, or the cross fins 15 can be used without the inner nozzle 13 and extend between the outer nozzle 12 to provide heat transfer from the combustion gases flowing to the liquid passing through the cross fins 15. also, in another embodiment of the cross fins 15, the cross fins 15 could have slots or openings formed across the two side of the fins in order to equalize the pressures acting on the fins to limit flexing of the fins during engine operation. The fluid passages through the cross fins 15 would be closed off from these openings to equalize pressures.
Operation of the nozzle assembly for the expander cycle rocket engine of the present invention is described below. One of the cryogenic liquids used for combustion (typically the hydrogen in a hydrogen oxygen combustion because the liquid hydrogen has the highest heat transfer coefficient of the two liquids) is passed through the primary or outer nozzle from bottom to top in order to cool the primary nozzle and heat up the liquid hydrogen. Some of the liquid hydrogen is also passed through the lower struts and into the inner nozzle and flows from bottom to top in order to also cool the inner nozzle and heat up the liquid hydrogen passing through. The heated up hydrogen then flows out from the inner nozzle through the upper struts to be joined with the heated up hydrogen leaving the primary or outer nozzle. The heated up hydrogen from both nozzles then flows through one or both of the turbo-pumps to pressurize one or both of the fuel and oxidizer for delivery to the combustion chamber within the nozzle assembly.
Use of the additional nozzle in the expander cycle rocket engine increases the available surface area of the nozzle assembly to provide the additional heat transfer necessary to drive the turbo-pumps for a higher mass flow required for higher pressures and higher thrust than is available using the prior art nozzle. The internal nozzle heat exchanger is substantially similar to the tubular heat exchanger used to cool the primary or outer nozzle. Its diameter and length is based on the magnitude of the additional heat transfer required. A plurality of struts attaches the internal heat exchange nozzle to the outer primary nozzle. The liquid propellant enters and exits the internal heat exchanger through ports contained within the struts. Multiple concentric internal heat exchangers could be used if required by engine sizing in other embodiments. Alternate embodiments using non-concentric, asymmetric or other geometries are also considered to be within the possible configurations of the present invention. In each embodiment, the primary flow path of the nozzle is unaffected. The envelope of the nozzle is adjusted as necessary to maintain the desired cross-sectional area of the primary flow path within the nozzle. However, the overall size, length and area ratio remain optimum. The internal heat exchanger elements are oriented along streamlines of the flow and the struts are aerodynamically shaped to mitigate any flow disturbances. The present invention removes the current thrust limitations of an expander cycle rocket engine and allows for engine designs of any thrust.
In an expander cycle rocket engine, liquid hydrogen is typically the fuel and liquid oxygen is typically the oxidizer. In this case, the liquid hydrogen is the fluid that is passed through the heat exchangers to cool the nozzles and drive the turbo-pump because the hydrogen has a greater coefficient of heat than does the oxygen. Because of the twin nozzle assembly of the present invention, a different fuel could be used such as kerosene, which is a much less volatile fuel than hydrogen. With kerosene as the fuel, liquid oxygen would be the oxidizer. In this case, the oxygen would have the higher heat capacity that the kerosene. As such, the liquid oxygen would be the fluid passed through the heat exchangers and used to drive the turbo-pump. The present invention is independent of the propellants selected for use in the engine.
