ENGINE WITH AFTERBURNER COMBUSTOR

CROSS-REFERENCE TO RELATED APPLICATIONS

[0001] The present disclosure claims priority to United States Provisional Patent Application No. 62/049,855 filed September 12, 2014; United States Provisional Patent Application No. 62/062,094 filed October 9, 2014; and United States Provisional Patent Application No. 62/085,538 filed November 29, 2014.

BACKGROUND

[0002] Numerous types of liquid propellant rocket engines are known. One type of engine utilizes an expander cycle in which pressurized fuel is expanded through a turbine prior to injection into a rocket combustion chamber. The turbine drives a fuel and/or oxidizer pump. Another type of engine utilizes a gas generator cycle in which fuel and oxidizer are burned in a pre-burner to generate an exhaust gas that is expanded through a turbine. The exhaust from the turbine is then dumped overboard. A third type of engine utilizes a staged combustion cycle that is similar to the gas generator cycle but the exhaust from the turbine is injected into the rocket combustion chamber rather than being dumped overboard.

SUMMARY

[0003] A liquid propellant rocket engine according to an example of the present disclosure includes a combustion chamber that has a throat and a nozzle that diverges from the throat. The nozzle has an afterburner combustor section that includes a fuel injector orifice. There is least one turbine and each such turbine includes a turbine inlet and a turbine outlet. The turbine inlet is fluidly coupled with an exhaust duct of a gas generator and the turbine outlet is fluidly coupled with the fuel injector orifice.

[0004] A further embodiment of any of the foregoing embodiments includes an oxidizer source fluidly coupled with an oxidizer injector orifice of the afterburner combustor section.

[0005] In a further embodiment of any of the forgoing embodiments, the gas generator is configured to produce a fuel-rich exhaust.

[0006] In a further embodiment of any of the forgoing embodiments, the nozzle includes a proximal half and a distal half with respect to proximity to the throat, and the afterburner combustor section is in the proximal half. [0007] A liquid propellant rocket engine according to an example of the present disclosure includes first and second combustion chambers having, respectively, first and second throats, and first and second nozzles that diverge, respectively, from the first and second throats. A fuel source is fluidly coupled with the first combustion chamber and a gas generator that has an exhaust duct. A turbine includes a turbine inlet and a turbine outlet. The turbine inlet is fluidly coupled with an exhaust duct of the gas generator and the turbine outlet is fluidly coupled with the second combustion chamber.

[0008] In a further embodiment of any of the forgoing embodiments, the first and second nozzles are of different sizes.

[0009] In a further embodiment of any of the forgoing embodiments, the turbine has a pressure ratio across the turbine inlet and the turbine outlet of approximately 2:1.

[0010] A method for a liquid propellant rocket engine according to an example of the present disclosure includes delivering fuel and oxidizer to a first combustion chamber that includes a first throat. The fuel and oxidizer react to generate a gas stream in a first nozzle that diverges from the first throat. Fuel and oxidizer are also delivered to a gas generator that includes an exhaust duct. The fuel and oxidizer react to generate a fuel-rich gas stream in the exhaust duct that expands across a turbine. The expanded fuel-rich gas stream from the turbine is delivered to one of a fuel injector orifice in an afterburner combustor section of the first nozzle, or a second, different combustion chamber.

[0011] In a further embodiment of any of the forgoing embodiments, the fuel-rich gas stream is expanded to a pressure that is substantially equal to or greater than 50% of the pressure of the gas stream in the first combustion chamber.

BRIEF DESCRIPTION OF THE DRAWINGS

[0012] The various features and advantages of the present disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows.

[0013] Figure 1 illustrates an example liquid propellant rocket engine with a gas generator and afterburner combustor section downstream of the main combustion chamber.

[0014] Figure 2 illustrates another example liquid propellant rocket engine that is similar to the engine in Figure 1 but additionally includes an additional turbine and oxidizer pump.

[0015] Figure 3 illustrates another liquid propellant rocket engine that includes a gas generator and an oxidizer source feeding the afterburner combustor. DETAILED DESCRIPTION

[0016] Figure 1 schematically illustrates selected portions of an example liquid propellant rocket engine 20. The engine 20 is mounted on a vehicle V (represented schematically) and, as will be described in further detail herein, is configured as an afterburning, gas generator cycle engine. The term "afterburning" and variations thereof refer to a burning of propellants to augment thrust generated by a main or primary combustion cycle.

[0017] The engine 20 includes a combustion chamber 22 that has a throat 24, which is a relatively narrow or narrowest portion of the combustion chamber 22. A nozzle 26 diverges from the throat 24 and includes an afterburner combustor section 28. In this example, the nozzle 26 includes a proximal section 26a and a distal section 26b with respect to proximity to the throat 24. The proximal and distal section 26a/26b are taken with regard the axial length of the nozzle 26 along central axis A. A line L in Figure 1 denotes the boundary between the proximal and distal sections 26a/26b. In one example, the proximal section 26a is approximately 10% of the axial length of the nozzle 26.

[0018] The nozzle 26 includes the afterburner combustor section 28. In this example, the afterburner combustor section 28 is in the proximal section 26a of the nozzle 26. In further examples, the afterburner combustor section 28 can be nearer (axially) to the throat 24 than to line L. In other examples, the afterburner combustor section 28 may be in the distal section 26b, although there may be greater performance gains if located in the proximal section 26a.

[0019] The afterburner combustor section 28 is generally an annular or frustoconical portion of the nozzle 26. For example, the afterburner combustor section 28 may include the portion of the wall of the nozzle 26 that includes or encompasses one or more fuel injector orifices 30 and one or more oxidizer injector orifices 32. The relative locations of the fuel injector orifices 30 with respect to the locations of the oxidizer injector orifices 32 (as seen in Figure 1) can be varied, in addition to variation in the type of injection provided by the orifices 30/32 with regard to injecting streams, fans, or sprays. In general though, the orifices 30/32 will be distributed circumferentially. In further examples, the fuel injector orifices 30 may be either forward or aft of the oxidizer injector orifices 32. Additionally, the orifices 30/32 may be configured to generate impinging injection flows. The orifices 30/32 additionally or alternatively can be configured to inject in a direction that is perpendicular to the central axis A, at an angle conformal to the inside surface of the nozzle 26, or at any angle in between. A further example of one or more of the orifices 30/32 can include an annular orifice.

[0020] The engine 20 also includes a gas generator 34. The gas generator 34 includes an exhaust duct 34a that is fluidly coupled with at least one turbine 36. The turbine 36 includes a turbine inlet 36a and a turbine outlet 36b. The turbine inlet 36a is fluidly coupled with the exhaust duct 34a and the turbine outlet 36b is fluidly coupled with the fuel injector orifices 30.

[0021] In this example, the engine 20 includes at least one oxidizer source 38 and at least one fuel source 40. A typical fuel can include, but is not limited to, kerosene, methane, or hydrogen. A typical oxidizer can include, but is not limited to, gaseous or liquid oxygen, nitrogen tetroxide, nitrous oxide, and hydrogen peroxide.

[0022] For example, the oxidizer source 38 may include one or more oxidizer pumps 38a, and the fuel source 40 may include one or more fuel pumps 40a, which are both mounted to be driven by the turbine 36. The oxidizer pump 38a is fluidly coupled with the combustion chamber 22, through primary oxidizer injectors 23a, as well as the secondary (afterburner) oxidizer injector orifices 32 and the gas generator 34. The fuel pump 40a is fluidly coupled with the gas generator 34 and the nozzle 26. Prior to injection into the combustion chamber 22, the fuel from the fuel pump 40a may be conveyed through internal passages in the walls of the nozzle 26, throat 24, and combustion chamber 22 for cooling, and to the primary fuel injector 22a.

[0023] During operation of the engine 20, the oxidizer source 38 delivers oxidizer and the fuel pump 40a delivers fuel. A portion of the fuel is delivered to the internal passages in the nozzle 26 as a coolant, and another portion of the fuel is delivered to the gas generator 34. The oxidizer is divided among the gas generator 34, the primary oxidizer injectors 23a for injection into the primary combustion chamber 22 and secondary injector orifices 32. The fuel and oxidizer burn in the gas generator 34 to generate a fuel-rich exhaust that is expanded through the one or more turbines 36. The expanded fuel-rich exhaust is then delivered from the turbine outlet 36b to the fuel injector orifices 30 in the afterburner combustor section 28. Fuel and oxidizer are also injected through injectors 22a into the combustion chamber 22 to generate thrust through the throat 24 and nozzle 26. The fuel-rich gas stream from the turbine 36 is injected through the fuel injector orifices 30 and oxidant is injected through the oxidant orifices 32 into the afterburner combustor section 28. The fuel-rich gas stream and oxidant burn downstream of the throat 24 to generate additional thrust in the nozzle 26. [0024] For enhanced afterburner thrust performance gains, the engine 20 may be controlled with respect to certain pressures in the system. For example, the combustion chamber 22 is configured to produce a first pressure therein and the gas generator 34 and turbine 36 are configured to produce a second pressure at the turbine outlet 36b that is close to the first pressure. By controlling the second pressure to be as closer to the first pressure than typical of a conventional Gas Generator Cycle Engine, greater augmentation thrust can be produced. The closer the second pressure is to the first pressure (low turbine pressure ratio), more fuel rich gases are required to drive the pumps and more oxidizer is required for afterburning to increase thrust augmentation.

[0025] The size of the turbine 36 can be selected with respect to a pressure ratio across the turbine inlet 36a and a turbine outlet 36b. For instance, if such a ratio is too high, such as 10: 1, the pressure at the turbine outlet 36b would be about 10% of first pressure in the combustion chamber 22, reducing the required gas generator flow and thrust efficiency. However, the use of a pressure ratio of approximately 2: 1 promotes greater thrust efficiency of the afterburner gases.

[0026] Figure 2 illustrates another example liquid propellant rocket engine 120 that is similar to the engine 20 but includes an oxidizer source 138 that has a first oxidizer pump 138a and a second oxidizer pump 138b. In this example, the first oxidizer pump 138a is fluidly coupled with the injectors 22a of the combustion chamber 22 and also the gas generator 34.

[0027] The engine 120 also includes a plurality of turbines 136, which in this example includes a first turbine 136-1 and second turbine 136-2. The first turbine 136-1 includes a turbine inlet 136a-l and a turbine outlet 136b-l, and the second turbine 136-2 includes a turbine inlet 136a-2 and a turbine outlet 136b-2. The first turbine 136-1 is coupled to, and drives, the fuel pump 40a and the first oxidizer pump 138a. The second turbine 136-2 is coupled to, and drives, the second oxidizer pump 138b.

[0028] Similar to the engine 20, the fuel-rich exhaust gas stream is expanded through the first turbine 136-1, but in this example is then also expanded through the second turbine 136-2 to drive the second oxidizer pump 138. The second oxidizer pump 138 serves to deliver the oxidizer to the oxidizer injector orifices 32.

[0029] Figure 3 illustrates another example liquid propellant rocket engine 220. In this example, rather than afterburning the fuel-rich gas stream from the one or more turbines 36, a second, different combustion chamber 50 serves for the afterburning. The second combustion chamber 50 includes a second throat 52 and a second nozzle 54 that diverges from the second throat 52. As can be appreciated, the first and second combustions chambers 22/50, first and second throats 24/52, and first and second nozzles 26/54 are separate and distinct from each other and thus discharge separate or uncommon effluent streams. For example, the first and second nozzles 26/54 are of different sizes, with the nozzle 54 being generally smaller. Likewise, the first and second combustions chambers 22/50 and the first and second throats 24/52 may also be of different sizes than each other.

[0030] In this example, the oxidizer pump 38 delivers oxidizer to the gas generator 34, the combustion chamber 22, and the combustion chamber 50. The fuel-rich gas stream from the one or more turbines 36 is delivered, in addition to oxidizer, into the second combustion chamber 50, which serves as an afterburner combustor, to provide afterburning thrust to augment the thrust generated by the main or primary first combustion chamber 22. Similar to the example above, the gas generator 34 and the one or more turbines 36 can be configured to provide a pressure at the turbine outlet 36b that is equal to or greater than 50% of the pressure in the combustion chamber 22.

[0031] The above examples and figures also embody exemplary methods for a liquid propellant rocket engine. For example, the method includes delivering fuel and oxidizer to the (first) combustion chamber 22, delivering fuel and oxidizer to the gas generator 34, expanding the fuel-rich gas stream across the turbine 36/136-1/136-2 and, depending on which of the engines 20/120/220, delivering the expanded fuel-rich gas stream from the turbine to one of: (i) the fuel injector orifices 30, or (ii) the second combustion chamber 50. In a further example, the method includes expanding the fuel-rich gas stream to a pressure that is substantially equal to approximately 50% of the pressure of the gas stream in the (first) combustion chamber 22.

[0032] Although a combination of features is shown in the illustrated examples, not all of them need to be combined to realize the benefits of various embodiments of this disclosure. In other words, a system designed according to an embodiment of this disclosure will not necessarily include all of the features shown in any one of the Figures or all of the portions schematically shown in the Figures. Moreover, selected features of one example embodiment may be combined with selected features of other example embodiments.

[0033] The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from this disclosure. The scope of legal protection given to this disclosure can only be determined by studying the following claims.