Film cooling in rotating detonation engines for secondary combustion.

Abstract

A rocket engine system having a coolant source including a coolant fuel, the coolant source for supplying the coolant fuel to a first surface of a wall that partially defines a combustion chamber of the rocket engine, the coolant fuel film cooling the first surface of the wall, and the oxidizer source for supplying the oxidizer to a second surface of the wall. The monitor is configured to control the flow of the coolant fuel and the oxidizer and to ensure that a stoichiometric ratio of the coolant fuel and oxidizer mixture is suitable for causing combustion of the coolant fuel and oxidizer mixture after the coolant fuel has film cooled the first surface of the wall.

Description

[Technical Field]

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS This application claims priority to and benefit of co-pending U.S. patent application Ser. No. 18 / 178,456 (Attorney Docket No. VENU-010), filed Mar. 3, 2023 by Andrew Thomas Duggleby and entitled "FILM COOLING WITH ROTATING DETONATION ENGINE TO SECONDARY COMBUSTION," and assigned to the assignee of the present application, and is hereby incorporated by reference in its entirety. [Background technology]

[0002] FIELD OF THE INVENTION Embodiments of the present invention generally relate to rockets, rocket engines, and their cooling systems. Summary of the Invention [Means for solving the problem]

[0003] Various embodiments of the present invention relate to rocket engine systems with improved cooling.

[0004] In one embodiment, the rocket engine and cooling system include a coolant source, a propellant source, a pressurization system, and a heat exchanger for supplying the coolant. In some embodiments, one or more of the coolant source and the propellant source are in operative communication with the pressurization system and the rocket engine so that the coolant can be pressurized and then heated by the heat exchanger. In various embodiments, the coolant is heated to a temperature and pressure such that the coolant is in a supercritical state. In various embodiments, the coolant is heated to a temperature and pressure below the temperature or pressure at which the coolant reaches the supercritical state. In various embodiments, the coolant is heated to a temperature and pressure above the temperature or pressure at which the coolant reaches the supercritical state.

[0005] In one embodiment, the propellant source includes one or more of a fuel, an oxidizer, and a coolant. In various embodiments, the oxidizer may be premixed with the fuel and coolant.

[0006] In various embodiments, the coolant may be a fuel, an oxidant, or an inert coolant.

[0007] In various other embodiments, the rocket engine and cooling system includes a cooling system with a coolant source for supplying coolant, a fuel system with a fuel source for supplying fuel, an oxidizer system with an oxidizer source for supplying oxidizer, a propellant pressurization system with a pressurization source for pressurizing the propellant, and a heat exchanger. In various embodiments, the pressurization source is in communication with the coolant after it has passed through the rocket engine and heat exchanger. Such an embodiment is referred to herein as an expander cycle.

[0008] In some embodiments, the improved rocket engine system includes a cooling system with a coolant source for supplying coolant, a fuel system with a fuel source for supplying fuel, an oxidizer system with an oxidizer source for supplying oxidizer, a propellant pressurization system with a pressurization source for pressurizing the propellant, and a heat exchanger. In some embodiments, the pressurization source is in communication with the coolant after it passes through the rocket engine and the heat exchanger (e.g., an expander cycle), and the aerospike nozzle is cooled by the coolant after the coolant powers the pressurization system.

[0009] Some embodiments of the present invention include a cooling system with a coolant source for supplying coolant, a fuel system with a fuel source for supplying fuel, an oxidizer system with an oxidizer source for supplying oxidizer, a propellant pressurization system with a propellant pressurization source for pressurizing the propellant, and a preburner. In one such embodiment, the pressurization source is driven by the coolant after it has passed through a rocket engine and a heat exchanger (e.g., an expander cycle). The preburner is used to heat the coolant to a temperature and pressure at which the coolant reaches a supercritical state by a secondary combustion reaction between the fuel and the oxidizer. In various embodiments, the preburner is used to heat the coolant to a temperature and pressure below the temperature or pressure at which the coolant reaches a supercritical state by a secondary combustion reaction between the fuel and the oxidizer. In various embodiments, the preburner is used to heat the coolant to a temperature and pressure above the temperature or pressure at which the coolant reaches a supercritical state by a secondary combustion reaction between the fuel and the oxidizer.

[0010] Some embodiments of the present invention include a cooling system that utilizes a coolant fuel for film cooling and then uses the coolant fuel to create secondary combustion. In one such embodiment, an oxidizer is mixed with the coolant fuel after the coolant fuel is used for film cooling. In various embodiments, the amount of coolant fuel and the amount of oxidizer are controlled to achieve a stoichiometric ratio of coolant fuel to oxidizer suitable for secondary combustion. [Brief explanation of the drawings]

[0011] These and other features of the present invention will be more fully understood by reference to the following detailed description taken in conjunction with the accompanying drawings.

[0012] [Figure 1]1 is a detailed schematic diagram of an embodiment of an improved rocket engine system that heats a coolant to a temperature and pressure at which the coolant becomes supercritical, or to a temperature or pressure below the temperature or pressure at which the coolant becomes supercritical, or to a temperature and pressure above the temperature or pressure at which the coolant becomes supercritical, in accordance with an embodiment of the present invention. FIG. 1 schematically illustrates a coolant source, a fuel source, an oxidizer source, a pressurization system, a heat exchanger, and a rocket engine, in accordance with an embodiment of the present invention.

[0013] [Figure 2] 1 is a detailed schematic diagram of an embodiment of an improved rocket engine system that heats a coolant to a temperature and pressure at which the coolant becomes supercritical, or to a temperature and pressure below or above the temperature and pressure at which the coolant becomes supercritical, in accordance with an embodiment of the present invention. FIG. 2 is a schematic diagram of a rocket engine with a coolant source, a fuel source, an oxidizer source, a pressurization system, a heat exchanger, and an aerospike nozzle, in accordance with an embodiment of the present invention.

[0014] [Figure 3] 1 is a detailed schematic diagram of an embodiment of an improved rocket engine system that heats a coolant to a temperature and pressure at which the coolant becomes supercritical, or to a temperature and pressure below or above the temperature and pressure at which the coolant becomes supercritical, in accordance with an embodiment of the present invention. FIG. 3 is a schematic diagram of a rocket engine with a coolant source, a fuel source, an oxidizer source, a pressurization system, a preburner, and an aerospike nozzle, in accordance with an embodiment of the present invention.

[0015] [Figure 4]4 is a detailed schematic diagram of an embodiment of an improved rocket engine system that heats a coolant to a temperature and pressure at which the coolant becomes supercritical, or to a temperature and pressure below or above the temperature and pressure at which the coolant becomes supercritical, in accordance with an embodiment of the present invention. FIG. 4 is a schematic diagram of a rocket engine with a coolant source, a fuel source, an oxidizer source, a pressurization system, an alternative preburner configuration, and an aerospike nozzle, in accordance with an embodiment of the present invention.

[0016] [Figure 5] 1 is a cross-sectional schematic diagram of an embodiment of an improved rocket engine system that applies a coolant to the rocket engine to film cool the rocket engine in accordance with an embodiment of the present invention. FIG.

[0017] [Figure 6] 1 is a perspective view of a continuous injection port and a slot injection port used to film cool a rocket engine by applying a coolant, according to an embodiment of the present invention.

[0018] [Figure 7] 1 is a cross-sectional schematic diagram of a rotary detonation engine according to an embodiment of the present invention.

[0019] [Figure 8] 1 is a cross-sectional view of a wall defining an annulus where detonation occurs in a rotary detonation engine and where one wall is treated, according to an embodiment of the present invention. FIG.

[0020] [Figure 9] 1 is a cross-sectional view of a wall defining an annulus where detonation occurs in a rotary detonation engine and where two walls are treated, according to an embodiment of the present invention. FIG.

[0021] [Figure 10]FIG. 1 is a schematic diagram of components of an improved rocket engine system that uses a coolant fuel for film cooling and then mixes the coolant fuel with an oxidizer for post-combustion, according to an embodiment of the present invention.

[0022] [Figure 11] 1 is a cross-sectional view of a wall defining a chamber in which combustion occurs in a rocket engine and in which the wall is treated, according to an embodiment of the present invention. FIG.

[0023] Unless otherwise noted, the drawings referred to in the brief description of the drawings should be understood as not being drawn to scale. DETAILED DESCRIPTION OF THE INVENTION

[0024] Various embodiments of the subject matter are described in detail below, examples of which are illustrated in the accompanying drawings. While various embodiments are discussed herein, it should be understood that they are not intended to be limited to these embodiments. Rather, the presented embodiments are intended to encompass alternatives, modifications, and equivalents, which may be included within the spirit and scope of the various embodiments. Furthermore, in describing the embodiments, numerous specific details are set forth in order to provide a thorough understanding of the embodiments of the subject matter. However, embodiments may be practiced without these specific details. In other instances, well-known methods, procedures, components, and circuits have not been described in detail so as not to unnecessarily obscure aspects of the described embodiments.

[0025] Throughout this specification, references to "one embodiment," "an embodiment," "embodiments," "various embodiments," "some embodiments," or similar terms mean that a particular feature, structure, or characteristic described in connection with that embodiment is included in at least one embodiment. Thus, the appearances of such phrases in various places throughout this specification do not necessarily all refer to the same embodiment. Furthermore, a particular feature, structure, or characteristic of any embodiment may be combined in any suitable manner with one or more other features, structures, or characteristics of one or more other embodiments, without limitation.

[0026] Although one or more embodiments of the present invention have been described in some detail for clarity of understanding, it will be apparent that certain changes and modifications may be practiced within the scope of the various embodiments. Accordingly, the described embodiments should be considered illustrative and not restrictive, and the scope of the various embodiments is not limited to the details set forth herein, but may be modified. In the various embodiments, none of the components and / or described steps suggest a specific order of operation unless explicitly stated therein.

[0027] It should be noted that the combustion process is essential to many types of combustion engines, including but not limited to rocket engines. The combustion process is usually defined as either "deflagration" or "detonation."

[0028] In deflagration combustion, the maximum velocity of the generated combustion wave is typically significantly less than the speed of sound. Furthermore, the blast overpressure generated in deflagration combustion is typically significantly less than 20 times the initial combustion pressure. Therefore, the deflagration combustion process is sometimes referred to as a nearly constant-pressure combustion.

[0029] In detonation combustion, flame speeds typically reach supersonic speeds and can, in fact, reach speeds on the order of thousands of meters per second. Additionally, the detonation combustion process can generate overpressures that can significantly exceed the initial combustion pressure by 20 to 100 times. Therefore, the detonation combustion process is sometimes referred to as a near-constant-volume combustion. Furthermore, compared to deflagration, detonation offers faster heat release, a greater entropy reduction, and greater thermal efficiency.

[0030] The advantageous properties of detonation combustion have led to the development of detonation-based propulsion engines (e.g., detonation-based rocket engines). One type of detonation-based rocket engine is the rotary detonation engine (RDE), sometimes referred to as a continuous rotary detonation engine (CRDE). For purposes of brevity and clarity, when describing various embodiments of the present invention, the following description may refer to rotary detonation engines, rotary detonation rocket engines, and the like. However, it should be noted that various embodiments of the present invention may also be well suited for use with various other types of detonation-based propulsion engines.

[0031] Continuing with reference to Figures 1-4, in various embodiments, the present invention is an improved rocket engine system that heats a coolant to a temperature and pressure at which the coolant becomes supercritical, or to a temperature and pressure below the temperature or pressure at which the coolant becomes supercritical, or to a temperature and pressure above the temperature or pressure at which the coolant becomes supercritical, in accordance with embodiments of the present invention. Those skilled in the art will appreciate that the flow circuits depicted in the various figures have been simplified so as not to obscure the invention with unnecessary detail. Additionally, those skilled in the art will appreciate that several valves, auxiliary lines, and bypass paths are present but may not be shown so as not to obscure the invention with unnecessary detail.

[0032] In various embodiments of the present invention, the coolant is composed entirely or partially of a non-reactive material, such as, but not limited to, carbon dioxide (CO), nitrogen (N), or water (H2O). In various embodiments of the present invention, the coolant is composed entirely or partially of a reactive material, such as, but not limited to, peroxide (H2O2), nitrous oxide (N2O), ammonia (NH3), or propane (C3H8).

[0033] In various embodiments of the present invention, the coolant may be composed entirely or partially of water, which provides several advantages. For example, when water is used as the coolant, the presence of water advantageously slows or retards the rate of reaction between the fuel and oxidizer, thereby improving rocket engine performance. This advantage is particularly important in detonation rocket engines, where a slower reaction rate allows for more efficient mixing of the fuel, oxidizer, etc., and subsequent detonation. Additionally, because water is denser than most conventional coolants, a given mass of water can be stored in a tank that is smaller (and therefore lighter) than would be required to store the same mass of a less dense conventional coolant. Furthermore, the higher density of water compared to conventional coolants allows the use of water to achieve a higher specific impulse (specific impulse is I sp Additionally, compared to the availability, toxicity, and cost of many traditional coolants, water is much more available, much less toxic, and much cheaper.

[0034] It should be noted that water can be used as a coolant in various embodiments of the present invention described below. Additionally, it should be noted that the following description of various embodiments of the present invention is well suited to, but not limited to, the various non-reactive materials and / or reactive coolants listed above.

[0035] Continuing with reference to FIG. 1 , in various embodiments of the present invention, a rocket engine system uses a propellant including a fuel source stored in a structure, such as a vehicle, connected to the rocket engine. The fuel is delivered to the engine through a fuel feed line 23. An oxidizer source, also stored in a structure, such as a vehicle, connected to the rocket engine, delivers oxidizer to the engine through an oxidizer feed line 22. FIG. 1 also schematically illustrates a coolant source stored in a structure, such as a vehicle, connected to the rocket engine. Coolant is delivered to the engine through a coolant feed line 19. The coolant source is in communication with a pressurization system. In various embodiments, the present invention includes a turbine 15, a coolant pump 16, a fuel pump 17, and an oxidizer pump 18. In various embodiments, the coolant pump 16 is in communication with a heat exchanger 11, such as through a high-pressure coolant line 9. In various embodiments, the fuel pump 17 is in communication with an injector manifold 10, such as through a high-pressure fuel line 7. In various embodiments, an oxidizer pump 18 communicates with the injector manifold 10, such as through a high pressure oxidizer line 8.

[0036] In various embodiments of the present invention, the temperature of the refrigerant is increased in heat exchanger 11 to a temperature and pressure at which the refrigerant becomes supercritical (hereinafter referred to as a supercritical refrigerant). In various embodiments of the present invention, the temperature of the refrigerant is increased in heat exchanger 11 to a temperature and pressure above the temperature or pressure at which the refrigerant becomes supercritical (hereinafter referred to as an above-supercritical refrigerant). In various embodiments of the present invention, the temperature of the refrigerant is increased in heat exchanger 11 to a temperature and pressure below the temperature or pressure at which the refrigerant becomes supercritical (hereinafter referred to as a sub-supercritical refrigerant).

[0037] In various embodiments of the present invention, the supercritical coolant then communicates with coolant channels integrated into the outer wall 4, such as through a coolant heat exchanger outlet line 12. In one embodiment, the supercritical state is a temperature and pressure just entering the supercritical range of the coolant used. For example, if water is used as the supercritical coolant, the temperature may be raised to 374-392°C and the pressure may be raised to 220-231 bar. Thus, the coolant may be raised to a just-supercritical state, just above the critical pressure and temperature, where the fluid has lower viscosity and higher conductivity, significantly increasing convective heat transfer. Internal coolant channels are integrated into the wall through manifolds and passages, as is well known to those skilled in the art. The coolant cools the engine wall, including a portion of the nozzle 2 and the throat 6, before returning to the heat exchanger 11 through a hot coolant inlet 13. After exchanging heat with the incoming coolant, the coolant exits the heat exchanger 11 through a hot coolant heat exchanger outlet 14 and enters the coolant turbine 15. Once the coolant has powered the pressurized system, it enters the injector manifold 10 through the turbine exit line 20 and enters the combustion chamber 1 along with the fuel and propellant, and exits the rocket engine through the throat 6.

[0038] In another embodiment of the invention, the sub-supercritical coolant communicates with coolant channels integrated into the exterior wall 4, such as through a coolant heat exchanger outlet line 12. For example, if water is used as the sub-supercritical coolant, the temperature may be raised to below 374-392°C and / or the pressure below 220-231 bar, so that by the time the sub-supercritical coolant reaches the most critical point in the cooling passage (e.g., directly beside the detonation wave in the case of an RDRE), it is brought to a just-supercritical state, just above the critical pressure and temperature. At this state, the fluid has lower viscosity and higher conductivity, significantly enhancing convective heat transfer. Internal coolant channels are integrated into the wall through manifolds and passages, as is well known to those skilled in the art. The coolant cools the engine wall, including a portion of the nozzle 2 and the throat 6, before returning to the heat exchanger 11 through the hot coolant inlet 13. After exchanging heat with the incoming coolant, the coolant exits the heat exchanger 11 through the hot coolant heat exchanger outlet 14 and enters the coolant turbine 15. Once the coolant has powered the pressurization system, it enters the injector manifold 10 through the turbine outlet line 20, enters the combustion chamber 1 along with the fuel and propellant, and exits the rocket engine through the throat 6.

[0039] In other embodiments of the invention, the supercritical coolant communicates with coolant channels integrated into the outer wall 4, such as through a coolant heat exchanger outlet line 12. For example, if water is used as the sub-supercritical coolant, the temperature may be elevated to above 374-392°C and / or the pressure may be above 220-231 bar. In such embodiments, the coolant is supercritical. Internal coolant channels are integrated into the wall through manifolds and passages, as is well known to those skilled in the art. The coolant cools the engine walls, including a portion of the nozzle 2 and the throat 6, before returning to the heat exchanger 11 through the hot coolant inlet 13. After exchanging heat with the incoming coolant, the coolant exits the heat exchanger 11 through the hot coolant heat exchanger outlet 14 and enters the coolant turbine 15. After the coolant powers the pressurization system, it enters the injector manifold 10 through the turbine outlet line 20, enters the combustion chamber 1 with the fuel and propellant, and exits the rocket engine through the throat 6.

[0040] 2, in various embodiments, a rocket engine system includes an aerospike nozzle 24 in which combustion occurs within an annulus 3 surrounded by an inner cowl 5 and an outer cowl 1. In various embodiments, the aerospike nozzle 24 may be any altitude compensating nozzle, such as, but not limited to, a plug nozzle, an expanding nozzle, a single expansion ramp nozzle, a stepped nozzle, an expansion deflection nozzle, or an extending nozzle.

[0041] In various embodiments, for example, if a rocket engine system has an aerospike nozzle, the rocket engine is a rotary detonation rocket engine, and there is an elevated but localized heat load near the injection point, introducing sub-supercritical coolant into the localized heat load area of ​​the rocket engine enhances cooling of the rocket engine by heating the sub-supercritical coolant to a supercritical state by the localized heat load area.

[0042] In various embodiments of the present invention, the inner cowl 5 contains coolant channels 4, and the outer cowl 1 contains coolant channels 21. The coolant (supercritical, sub-supercritical, or super-supercritical coolant) from the heat exchanger outlet 12 first cools the inner cowl 5 through the coolant channels 4 before returning to the heat exchanger 11 as "hot coolant" through the hot coolant heat exchanger inlet 13. After exchanging heat with the incoming coolant, the hot coolant exits the heat exchanger 11 through the hot coolant heat exchanger outlet 14 and enters the coolant turbine 15. After the turbine 15, the coolant returns to the aerospike engine and cools the outer cowl 1 through the coolant channels 21. The coolant channels 4 and 21 are integrated into the cowls through manifolds and passages, as is well known to those skilled in the art. After the coolant powers the pressurization system, it enters the injector manifold 10 through the turbine outlet line 20, enters the combustion chamber annulus 3 along with the fuel and propellant, and exits the rocket engine through the throat 6.

[0043] 3, in various embodiments of the present invention, a rocket engine system uses a preburner 25 to add heat to a coolant (supercritical, sub-supercritical, or super-supercritical coolant) either completely or temporarily, for example, only during start-up, replacing or contributing to a heat exchanger. In various embodiments, fuel is diverted to the preburner from high-pressure fuel line 7 through fuel preburner inlet 26, and oxidizer is diverted to the preburner from high-pressure oxidizer line 8 through oxidizer preburner inlet 27.

[0044] Continuing with reference to FIG. 4, in various embodiments of the present invention, a rocket engine system uses a preburner 25 that powers a pressurization system, which is then mixed with a coolant (supercritical, sub-supercritical, or super-supercritical) in the preburner 25 to power the turbopump 15 through turbine inlet line 28 and then cool the rocket engine through engine coolant line 20.

[0045] Continuing with reference to FIG. 5, a cross-sectional schematic diagram 500 illustrating an embodiment of an improved rocket engine system is provided. As shown in FIG. 5, in accordance with an embodiment of the present invention, a coolant is applied to rocket engine 502 to film cool rocket engine 502. In one embodiment of the present invention, the coolant is water. More specifically, in one embodiment, water is applied to the interior surface of rocket engine 502. In various embodiments, applying water to the interior surface of rocket engine 502 provides both film cooling and a protective barrier to the interior surface of rocket engine 502. By providing a protective barrier, embodiments of the present invention result in a more reliable rocket engine. Specifically, the protective barrier provided by this embodiment of film cooling reduces the frequency of inspection of, and even the need to replace, the combustion wall chamber of rocket engine 502. Accordingly, embodiments of the present invention are well suited for use with reusable rocket engines. That is, various embodiments of the present invention may extend the life of a reusable rocket and / or increase the number of times a reusable rocket can be used.

[0046] Referring again to Figure 5, as noted above, because water is denser than most conventional coolants, various embodiments of the present invention allow a given mass of water to be stored in a tank that is smaller (and therefore lighter) than would be required to store the same mass of a less dense conventional coolant. Furthermore, because water is denser than many conventional coolants, embodiments of the present invention provide improved specific impulse compared to less dense conventional coolants (e.g., the fuel commonly used in all rocket engines currently employing film cooling). Additionally, compared to the availability, toxicity, and cost of many conventional coolants, the water coolant used in various embodiments of the present invention is more readily available, less toxic, and less expensive.

[0047] 5, in various embodiments, water is applied to the interior surfaces of rocket engine 502 at or near combustion chamber 506. As previously mentioned, when water is used as a coolant, its presence advantageously reduces or retards the rate of reaction between the fuel and oxidizer in combustion chamber 506, thereby improving the performance of rocket engine 502. This benefit is particularly important in detonation rocket engines, as the reduced reaction rate allows for more efficient mixing of the fuel, oxidizer, etc., and subsequent detonation in the combustion chamber.

[0048] Continuing to refer to FIG. 5, various ports 504a and 504b are located in or adjacent to combustion chamber 506. Ports 504c and 504d are also shown in FIG. 5. FIG. 6, discussed below, details an example port configuration that may be used in accordance with various embodiments of the present invention. Returning to FIG. 5, one or more of ports 504a, 504b, 504c, and 504d are used to apply water to the interior surface of rocket engine 502. Additionally, it should be noted that various embodiments of the present invention are well suited to having more or fewer ports. Various embodiments of the present invention are well suited to arranging ports in locations other than those illustrated in FIG. 5.

[0049] Referring again to Figure 5, it should also be noted that the film cooling embodiment corresponding to Figure 5 is well suited for use in combination with any of the interior wall cooling embodiments described in detail above and corresponding to Figures 1-4. Additionally, the film cooling embodiment of Figure 5 is also well suited for use in combination with rocket engine cooling techniques such as ablative layers.

[0050] While the above description of the embodiment of FIG. 5 (including embodiments used in combination with embodiments related to FIGS. 1-4) specifically describes the use of water as a coolant, these various embodiments are well suited for use with coolants composed entirely or in part of non-reactive materials other than water. Such non-reactive materials include, but are not limited to, carbon dioxide (CO), nitrous oxide (NO), or nitrogen (N). In addition, the various embodiments corresponding to FIG. 5 (including embodiments used in combination with embodiments related to FIGS. 1-4) are well suited for use with coolants composed entirely or in part of reactive materials. Such reactive materials include, but are not limited to, peroxide (HO), ammonia (NH), or propane (CH).

[0051] 5, in various embodiments of the invention, the water is at a temperature and pressure such that the water becomes supercritical when applied to the interior surface of rocket engine 502. In other various embodiments of the invention, the water is at a temperature and pressure above the temperature or pressure at which the water becomes supercritical when applied to the interior surface of rocket engine 502. Also, in various embodiments of the invention, the water is at a temperature and pressure below the temperature or pressure at which the coolant becomes supercritical when applied to the interior surface of rocket engine 502. In such embodiments of the invention, once the water is applied to the interior surface of the rocket engine, the temperature or pressure of the water is adjusted (i.e., increased or decreased) so that the temperature and pressure of the water corresponds to the supercritical state of water. For example, if water is introduced to the interior surface of rocket engine 502 in a sub- or supercritical state, at the point of maximum heat generation (e.g., adjacent the detonation wave in the case of an RDE), the interior surface of rocket engine 502 will regulate the temperature of the water to a range of approximately 374-392°C and / or regulate the pressure to a pressure of 220-231 bar. In such an embodiment, this regulation of the temperature and pressure of the water will result in a supercritical state, which significantly enhances convective heat transfer due to the lower viscosity and higher conductivity of water in the supercritical state.

[0052] Referring again to FIG. 5 , in various embodiments, for example, when a rocket engine system has an aerospike nozzle, the rocket engine is a rotary detonation rocket engine, and there is elevated but localized heat load near the injection point, embodiments of the present invention inject water into the localized hot area on the inner surface of the rocket engine 502 to enhance cooling of the rocket engine 502.

[0053] Continuing with reference to FIG. 6, a perspective view of a continuous injection port 602 and a slot injection port 604 are provided for use in applying a coolant (e.g., but not limited to, water) to the interior surface of a rocket engine 502 (FIG. 5), as described in connection with the embodiment of FIG. 5. As shown in FIG. 6, the continuous injection port 602 is configured to distribute a coolant, such as, but not limited to, water, as a substantially continuous flow onto the interior surface of the rocket engine 502. Arrows 606 illustrate the direction of coolant flow along the interior surface of the rocket engine 502. Similarly, in FIG. 6, the slot injection port 604 is configured to distribute a coolant, such as, but not limited to, water, as a discontinuous flow onto the interior surface of the rocket engine 502. Arrows 608 illustrate the direction of coolant flow along the interior surface of the rocket engine 502. While FIG. 6 illustrates continuous injection ports 602 and slotted injection ports 604, embodiments of the present invention are well suited for use with a variety of other features, ports, and port shapes, types, and configurations to enable the application of a coolant, such as, but not limited to, water, onto the interior surfaces of rocket engine 502.

[0054] The convective heat flux to the coolant, q = hΔT, is given by the convection coefficient, h, and the temperature difference, ΔT = T 燃焼 -T 冷却剤 In the supercritical state, the viscosity of the coolant decreases and the thermal conductivity increases, resulting in a large increase in the convection coefficient h. The coolant temperature T 冷却剤 Although this increases the heat transfer rate and reduces ΔT, it increases the total heat transfer rate, thus allowing the rocket engine to be cooled much more effectively and efficiently.

[0055] Continuing with reference to Figure 7, a cross-sectional schematic diagram of a rotary detonation engine 700 is provided. As will be described below, various embodiments of the present invention advantageously cool the rotary detonation engine 700 and provide secondary combustion to increase the thrust of the rotary detonation engine 700. It should be noted that Figure 7 does not show various well-known components and configurations in detail so as not to unnecessarily obscure aspects of the various embodiments of the present invention being described.

[0056] Figure 7 schematically illustrates an air inlet region 702 and a combustion chamber 704. It will be appreciated that the air inlet region 702, in some embodiments, can direct air into the combustion chamber 704. As shown in Figure 7, the combustion chamber 704 generally has annular walls, designated 706, that define an annular region 708 in which detonation occurs. For purposes of clarity, the walls 706 are depicted schematically in Figure 7; however, a more detailed depiction of these walls and a corresponding detailed description are provided below.

[0057] Additionally, as will be described in more detail below, in various embodiments of the invention, combustion chamber 704 may include multiple annuli and corresponding multiple walls defining the multiple annuli. In such embodiments, detonation may occur in one or more of the annuli. Furthermore, in embodiments of the invention having a combustion chamber comprised of multiple annuli, detonation may occur in multiple or even all of the annuli simultaneously.

[0058] Continuing with reference to FIG. 7 , region 710 is illustrated. Region 710 is the portion of rotary detonation engine 700 where coolant is supplied to wall 706, as described above in connection with the embodiment of FIGS. 1-6 . Note that some embodiments of the present invention will include one or more of the structures, features, and processes described in the embodiments corresponding to FIGS. 1-6 above. More specifically, some embodiments of the present invention will include turbine 15, coolant pump 16, fuel pump 17, and oxidizer pump 18, etc., all of which are shown in FIG. 1 . Furthermore, in various embodiments of the present invention, coolant pump 16 is in communication with heat exchanger 11, such as through high-pressure coolant line 9. Also, in various embodiments, fuel pump 17 is in communication with injector manifold 10, such as through high-pressure fuel line 7. In various embodiments, oxidizer pump 18 is in communication with injector manifold 10, such as through high-pressure oxidizer line 8. For purposes of brevity and clarity, some of the detailed descriptions of the structures, features, and processes of the embodiments described above in connection with FIGS. 1-6 will not be repeated in their entirety below.

[0059] Continuing with reference to Figure 8, a cross-sectional view 800 of wall 706 is provided, including region 710, both of which are shown in Figure 7. For purposes of clarity, because wall 706 is annular and Figure 8 is a cross-sectional view, an axis of symmetry 802 is shown to indicate the symmetrical orientation of the features depicted in Figure 8. Additionally, Figure 8 details wall 706 to clearly show that two portions 706a and 706b define annular portion 708 where detonation occurs.

[0060] In embodiments of the invention, fuel is applied to the inside of wall portion 706b (i.e., the side closest to annulus 708). More specifically, in various embodiments, fuel is supplied through ports 806 and channels 808, thereby providing film cooling along the inside of wall portion 706b via the supplied fuel. Embodiments of the invention are well suited for use with various other features, cooling channels, port shapes, port types, and port configurations to enable the application of coolant fuel onto the interior surface of wall portion 706b, for example, as described above in accordance with the detailed description of the embodiments of Figures 1-6. In Figure 8, the direction of fuel flow during film cooling is illustrated by line 804. In various embodiments, applying fuel to the inside of wall portion 706b provides film cooling and also provides a protective barrier to the inside of wall portion 706b. By providing a protective barrier, embodiments of the invention result in a more reliable rocket engine. In particular, the protective barrier provided by the fuel and corresponding film cooling on the inside of wall 706b reduces the frequency of inspecting wall 706b and even the need to replace it. Accordingly, embodiments of the present invention are well suited for use with reusable rocket engines. That is, various embodiments of the present invention may extend the life of a reusable rocket and / or increase the number of times a reusable rocket can be used.

[0061] Continuing with reference to FIG. 8 , various embodiments of the present invention also provide oxidizer along the exterior of wall 706b (i.e., the side furthest from annular portion 708). In FIG. 8 , the direction of oxidizer flow along the exterior surface of wall 706b is illustrated by line 810. More specifically, in various embodiments, oxidizer is provided through one or more ports (not shown). Furthermore, embodiments of the present invention are well suited for use with various other features, cooling channels, port shapes, port types, and port configurations to provide oxidizer along the exterior surface of wall 706b, as detailed above in accordance with the embodiments of FIGS. 1-6 . It should be noted that embodiments of the present invention are also well suited for use with various oxidizers, including, but not limited to, air, liquid oxygen, hydrogen peroxide, and the like. Furthermore, in various embodiments of the present invention, flowing oxidizer along the exterior surface of wall 706b, as illustrated by line 810, results in cooling of wall 706b.

[0062] Referring again to FIG. 8 , in various embodiments of the present invention, the oxidizer comprises at least a portion of air, which is introduced, such as at air inlet region 702 in FIG. 7 , and then directed along the outer surface of wall 706b, as illustrated by line 810. In various embodiments, the air may be introduced, such as at air inlet region 702, at subsonic or supersonic speeds. In some embodiments of the present invention, the flow rate of the air along line 810 is adjusted to achieve a desired air flow velocity. Additionally, in various embodiments of the present invention, the oxidizer may comprise multiple oxidizers. As one example, air introduced at supersonic speeds can actually generate heat when directed along the outer surface of wall 706b. In such cases, embodiments of the present invention may provide liquid oxygen or the like along with the introduced air to appropriately temperature-control the outer surface of wall 706b.

[0063] Continuing to refer to Figure 8, in an embodiment of the present invention, the fuel used for film cooling of the inner surface 706b and the oxidizer directed along the outer surface of the wall 706b will mix after passing the aft edge 812 of the wall 706b. Figure 8 shows a circle 814 that figuratively depicts the region where the coolant fuel and oxidizer mix. It should be noted that the circle 814 is merely a diagrammatic representation of the mixing of the coolant fuel and oxidizer, and that mixing of the coolant fuel and oxidizer may occur in regions other than within or near the circle 814.

[0064] 8, in various embodiments of the present invention, secondary combustion occurs when the fuel used to film cool the inner surface 706b and the oxidizer directed along the outer surface of the wall 706b mix together past the aft end 812 of the wall 706b. In embodiments of the present invention, secondary combustion occurs due to an ignition source readily present outside the annular portion 708.

[0065] Referring again to FIG. 8 , it can be seen that specific impulse (commonly abbreviated as Isp) is an important factor in determining the efficiency of thrust generated by a rotary detonation rocket engine, such as a rotary detonation rocket engine. More specifically, the Isp value measures how efficiently a reacting mass engine, such as the rotary detonation rocket engine 700 of FIG. 7 , generates thrust from a propellant, such as fuel. In embodiments of the present invention, additional thrust is generated for the rotary detonation rocket engine 700 through secondary combustion of a coolant fuel and an oxidizer. As a result, embodiments of the present invention improve the efficiency, i.e., Isp value, of the rotary detonation rocket engine 700. Additionally, embodiments of the present invention are advantageous and unique in that they achieve a “two-in-one” benefit by first using coolant fuel for film cooling and then using the same coolant fuel to generate secondary combustion and corresponding additional thrust. Therefore, unlike conventional film cooling approaches, embodiments of the present invention do not simply vent or waste the film coolant. Instead, embodiments of the present invention can cool surfaces within the annulus 708 without wasting payload, wasting propellant, or impeding the flow of propellant through the annulus 708. Additionally, embodiments of the present invention derive additional thrust from the coolant fuel via secondary combustion. As previously mentioned, for purposes of brevity and clarity, portions of this detailed description refer to rotary detonation engines, rotary detonation rocket engines, and the like when describing various embodiments of the present invention. However, it should be noted that various embodiments of the present invention are well suited for use with various other types of detonation-based propulsion engines. Furthermore, in various embodiments of the present invention, mixing of the coolant fuel with the oxidizer occurs only after the coolant fuel is no longer present within the annulus 708. That is, in embodiments of the present invention, the coolant fuel is utilized for film cooling within the annulus 708, and once the coolant fuel is no longer present within the annulus 708 (e.g., beyond the aft end 812 of the wall 706b), the coolant fuel is used for secondary combustion.

[0066] In embodiments of the present invention, the amount of coolant fuel and the amount of oxidizer are controlled to achieve a stoichiometric ratio of the coolant fuel and oxidizer suitable for combustion. More specifically, in embodiments of the present invention, the flow rate or volume of the coolant fuel and / or the flow rate or volume of the oxidizer are adjusted to achieve a stoichiometric ratio of the coolant fuel and oxidizer suitable for combustion upon mixing. As an example, when air is used as the oxidizer, as the altitude and / or speed of the rotary detonation engine 700 changes, the flow rate or volume of the coolant fuel and / or the flow rate or volume of the oxidizer are adjusted to achieve a stoichiometric ratio of the coolant fuel and oxidizer suitable for combustion upon mixing.

[0067] In one embodiment of the invention, if excess oxidizer is already present after detonation at the annulus 708, then only the flow rate or volume of coolant fuel may need to be adjusted to achieve a stoichiometry of the coolant fuel and oxidizer mixture (beyond the aft end 812 of the wall 706b) suitable for secondary combustion. Similarly, in one embodiment of the invention, if excess coolant fuel is already present after detonation at the annulus 708, then only the flow rate or volume of oxidizer may need to be adjusted to achieve a stoichiometry of the coolant fuel and oxidizer mixture (beyond the aft end 812 of the wall 706b) suitable for secondary combustion. As yet another example, if air is used as the oxidizer and there is insufficient flow rate or volume of air when mixed with the coolant fuel to achieve a combustible stoichiometry, then an embodiment of the invention would provide liquid oxygen, or the like, along with the entrained air to achieve the proper combustible stoichiometry when mixed beyond the aft end 812 of the wall 706b. Thus, various embodiments of the present invention differ from other film cooling approaches in that they adjust the coolant fuel flow rate or volume and / or adjust the oxidizer flow rate or volume to make the stoichiometric ratio of the coolant fuel and oxidizer mixture (beyond the aft end 812 of the wall 706b) suitable for secondary combustion.

[0068] Continuing to refer to Figure 8, embodiments of the present invention are well suited for use in combination with various other cooling techniques, including, but not limited to, those described in detail above in accordance with the embodiments of Figures 1-6. For example, in some embodiments of the present invention, wall portion 706a is cooled as described in the embodiments corresponding to Figures 1-6, and wall portion 706b is treated as described in the embodiments corresponding to Figures 7-8.

[0069] Continuing with reference to FIG. 9 , a cross-sectional view 900 of walls 706 a and 706 b is shown, where walls 706 a and 706 b are both as described in the embodiment corresponding to FIGS. 7-8 . It will be understood that in the embodiment of FIG. 9 , wall 706 b is treated as described in the embodiment corresponding to FIGS. 7-8 . For purposes of brevity and clarity, the discussion of wall 706 b and the embodiment corresponding to FIGS. 7-8 will not be repeated here. Furthermore, in this embodiment, wall 706 a is treated in the same or similar manner as described for wall 706 b in the embodiment corresponding to FIGS. 7-8 . Again, for purposes of brevity and clarity, the discussion of wall 706 a and the embodiment corresponding to FIG. 9 does not fully repeat the discussion corresponding to the embodiment of FIGS. 7-8 . However, it should be noted that various embodiment variations noted in the discussion corresponding to FIGS. 7-8 are also applicable to the embodiment of FIG. 9 . With respect to wall 706a, in embodiments of the invention, fuel is applied to the inside of wall 706a (i.e., the side closest to annulus 708). More specifically, in various embodiments, fuel is supplied through ports 906 and channels 908, thereby providing film cooling along the inside of wall 706a via the supplied fuel. Embodiments of the invention are well suited for use with various other features, cooling channels, port shapes, port types, and port configurations to enable the application of coolant fuel onto the interior surface of wall 706a, for example, as described above in accordance with the detailed description of the embodiments of Figures 1-6. In Figure 9, the direction of fuel flow during film cooling is illustrated by line 904. In various embodiments, applying fuel to the inside of wall 706a provides film cooling and a protective barrier to the inside of wall 706a.

[0070] Continuing with reference to FIG. 9 , various embodiments of the present invention also provide oxidizer along the exterior of wall 706a (i.e., the side furthest from annular portion 708). In FIG. 9 , the flow direction of oxidizer along the exterior surface of wall 706a is illustrated by line 910. More specifically, in various embodiments, oxidizer is provided through one or more ports (not shown). Furthermore, embodiments of the present invention are well suited for use with various other features, cooling channels, port shapes, port types, and port configurations to provide oxidizer along the exterior surface of wall 706a, as described in detail above in accordance with the embodiments of FIGS. 1-6 . It should be noted that embodiments of the present invention are also well suited for use with various oxidizers, including, but not limited to, air, liquid oxygen, hydrogen peroxide, and the like. Furthermore, in various embodiments of the present invention, flowing oxidizer along the exterior surface of wall 706a, as illustrated by line 910, results in cooling of wall 706a.

[0071] Continuing to refer to Figure 9, in an embodiment of the present invention, the fuel used for film cooling of the inner surface 706a and the oxidizer directed along the outer surface of the wall 706a mix after passing the aft end 912 of the wall 706a. In Figure 9, a circle 914 is shown that figuratively depicts the region where the coolant fuel and oxidizer mix. It should be noted that the circle 914 is merely a diagrammatic representation of the mixing of the coolant fuel and oxidizer, and that mixing of the coolant fuel and oxidizer may occur in regions other than within or near the circle 914.

[0072] Referring again to Figure 9, in various embodiments of the present invention, secondary combustion occurs when the fuel used to film cool the inner surface 706a and the oxidizer directed along the outer surface of the wall 706a mix past the aft end 912 of the wall 706a. In embodiments of the present invention, the secondary combustion occurs from an ignition source readily present outside of the annulus 708. Thus, in the embodiment of Figure 9, there are two secondary combustions. It should also be noted that embodiments of the present invention are well suited for use in detonation-based propulsion engines having multiple annulus sections.

[0073] Continuing with reference to Figure 10, a schematic diagram 1000 is provided illustrating various components of a system according to an embodiment of the present invention. The components of schematic diagram 1000 may be used in accordance with the embodiments described above, for example, in connection with Figures 7-9. Additionally, the components of schematic diagram 1000 may be used in accordance with the embodiments described below, for example, in connection with Figure 11.

[0074] 10 , in accordance with an embodiment of the present invention, a system is comprised of a fuel tank 1002 for supplying coolant fuel and an oxidizer tank 1004 for supplying oxidizer. Additionally, in an embodiment of the present invention, the system also includes a monitor 1006 connected to control valves 1008 and 1010. In various embodiments, the monitor 1006 utilizes the control valves 1008 and 1010 to adjust the amount of coolant fuel and the amount of oxidizer to the required stoichiometric ratio suitable for secondary combustion to occur, such as in the rotary detonation engine 700. In various embodiments, the present invention further includes a heat exchanger 1012 and a pressurized gas source 1014.

[0075] Continuing with reference to FIG. 11 , a cross-sectional view 1100 of a combustion chamber wall 1106 of a propulsion engine is shown. In various embodiments, fuel is applied to the inside of the wall 1106 (i.e., the side closest to the combustion zone 1108). More specifically, in various embodiments, fuel is supplied through ports 1110 and channels 1112, allowing film cooling along the inside of the wall 1106 via the supplied fuel. Embodiments of the present invention are well suited for use with various other features, cooling channels, port shapes, port types, and port configurations to allow coolant fuel to be applied onto the interior surface of the wall 1106, for example, as described above in accordance with the detailed description of the embodiments of FIGS. 1-6 . In FIG. 11 , the direction of coolant fuel flow during film cooling is illustrated by line 1104. In various embodiments, applying fuel to the inside of the wall 1106 allows the fuel to provide film cooling and a protective barrier to the inside of the wall 1106. By providing a protective barrier, embodiments of the present invention result in a more reliable rocket engine. In particular, the protective barrier provided by the fuel and corresponding film cooling on the inside of the wall 1106 reduces the frequency of inspecting and even the need to replace the wall 1106. Accordingly, embodiments of the present invention are well suited for use with reusable rocket engines. That is, various embodiments of the present invention may extend the life of a reusable rocket and / or increase the number of times a reusable rocket can be used.

[0076] Continuing with reference to FIG. 11 , various embodiments of the present invention also provide oxidizer along the exterior of wall 1106 (i.e., the side furthest from combustion chamber 1108). In FIG. 11 , the flow direction of oxidizer along the exterior surface of wall 1106 is illustrated by lines 1114 and 1116. More specifically, in various embodiments, oxidizer is provided through one or more ports (not shown). Furthermore, embodiments of the present invention are well suited for use with various other features, cooling channels, port shapes, port types, and port configurations to provide oxidizer along the exterior surface of wall 1106, as detailed above in accordance with the embodiments of FIGS. 1-6 . It should be noted that embodiments of the present invention are also well suited for use with various oxidizers, including, but not limited to, air, liquid oxygen, hydrogen peroxide, etc. Furthermore, in various embodiments of the present invention, flowing oxidizer along the exterior surface of wall 1106 as illustrated by lines 1114 and 1116 results in cooling of wall 1106.

[0077] 11 , in various embodiments of the present invention, the oxidizer is comprised at least in part of air, which is introduced, such as at an air inlet region of the rocket, and then directed along the exterior surface of wall 1106 as shown by lines 1114 and 1116. Additionally, in various embodiments of the present invention, the oxidizer may be comprised of multiple oxidizers. As detailed above, in various embodiments, the amount of coolant fuel and the amount of oxidizer are controlled to achieve a stoichiometric ratio of coolant fuel to oxidizer suitable for secondary combustion.

[0078] Continuing to refer to Figure 11, in an embodiment of the present invention, the fuel used for film cooling of the inner surface 1106 and the oxidizer directed along the outer surface of the wall 1106 mix after passing the aft ends 1106a and 1106b of the wall 1106. Circles 1118 and 1120 are shown in Figure 11 to depict the regions where the coolant fuel and oxidizer mix. It should be noted that circles 1118 and 1120 are merely a graphical representation of the mixing of the coolant fuel and oxidizer, and that mixing of the coolant fuel and oxidizer may occur in regions other than within or near circles 1118 and 1120.

[0079] Further, in this embodiment, wall 1106 is treated in the same or similar manner as described for wall 706b in the embodiment corresponding to Figures 7-9. Again, for purposes of brevity and clarity, the discussion of wall 1106 and the embodiment corresponding to Figure 11 does not fully repeat the discussion corresponding to the embodiment of Figures 7-9. However, it should be noted that various embodiment variations mentioned in the discussion corresponding to Figures 7-9 are also applicable to the embodiment of Figure 11.

[0080] 11, in various embodiments of the present invention, secondary combustion occurs when the fuel used to film cool the inner surface 1106 and the oxidizer directed along the outer surface of the wall 1106 mix together past the aft ends 1106a and 1106b of the wall 1106. In embodiments of the present invention, secondary combustion occurs due to an ignition source readily present outside the combustion chamber 1108.

[0081] The above description of the embodiments is not intended to be exhaustive or to limit the embodiments to the precise forms described. Rather, the example embodiments in the description of the embodiments are presented to enable one skilled in the art to make and use embodiments of the described subject matter. Also, while various embodiments are described in various combinations, any two or more embodiments may be combined. While some embodiments are described in language specific to structural features and / or methodological acts, it should be understood that the subject matter defined in the appended claims is not necessarily limited to the specific features or acts described above. Rather, the specific features and acts described above are disclosed by way of illustration and as example forms of implementing the claims and their equivalents.

[0082] Generally, the present description discloses at least the following:

[0083] A rocket engine system having a coolant source including a coolant fuel, the coolant source for supplying the coolant fuel to a first surface of a wall that partially defines a combustion chamber of the rocket engine, the coolant fuel film cooling the first surface of the wall, and an oxidizer source for supplying an oxidizer to a second surface of the wall, the monitor configured to control the flow of the coolant fuel and the flow of the oxidizer so that a stoichiometric ratio of the coolant fuel and oxidizer mixture is suitable for causing combustion of the coolant fuel and oxidizer mixture after the coolant fuel has film cooled the first surface of the wall.

[0084] The present description further discloses at least the following embodiments:

[0085] A first embodiment of the technology herein is

[0086] a coolant source including a coolant fuel configured to supply the coolant fuel to a first surface of a wall that partially defines an annulus in which detonation occurs in the rotary detonation engine, the coolant fuel film cooling the first surface of the wall;

[0087] an oxidant source configured to supply oxidant to a second surface of the wall that partially defines the annulus; and

[0088] The detonation engine system includes a monitor configured to control the flow of the coolant fuel and the flow of the oxidizer, the monitor being further configured to ensure that a stoichiometric ratio of the mixture of the coolant fuel and the oxidizer is suitable for causing combustion of the mixture of the coolant fuel and the oxidizer after the coolant fuel has completed film cooling the first surface of the wall.

[0089] A further embodiment of any of the preceding or following rocket engine system embodiments, wherein the coolant fuel film cools the first surface of the wall, thereby providing a protective barrier to the first surface of the wall.

[0090] A further embodiment of any of the preceding or following rocket engine system embodiments, comprising:

[0091] An embodiment further comprising at least one port positioned to apply the coolant fuel to a first surface of the wall.

[0092] A further embodiment of any of the preceding or following rocket engine system embodiments, wherein the oxidizer comprises liquid oxygen.

[0093] A further embodiment of any of the preceding or following rocket engine system embodiments, wherein the oxidizer comprises air.

[0094] Further embodiments of any of the preceding or following rocket engine system embodiments, wherein the oxidizer comprises hydrogen peroxide.

[0095] In a further embodiment of any of the preceding or following rocket engine system embodiments, the wall comprises:

[0096] further having at least one cooling passage formed therein;

[0097] An embodiment wherein the coolant source is fluidly connected to the at least cooling passage.

[0098] A further embodiment of any of the preceding or following rocket engine system embodiments, comprising:

[0099] a coolant source including a coolant fuel, the coolant source configured to supply the coolant fuel to a first surface of a wall that partially defines a combustion chamber of the rocket engine, the coolant fuel providing film cooling to the first surface of the wall;

[0100] an oxidizer source configured to supply oxidizer to a second surface of the wall that partially defines the combustion chamber; and

[0101] An embodiment including a monitor configured to control the flow of the coolant fuel and the flow of the oxidant, the monitor further configured to ensure that a stoichiometric ratio of the coolant fuel and the oxidant mixture is suitable for causing combustion of the coolant fuel and the oxidant mixture after the coolant fuel has completed film cooling the first surface of the wall.

[0102] A further embodiment of any of the preceding or following rocket engine system embodiments, wherein the coolant fuel film cools the first surface of the wall, thereby providing a protective barrier to the first surface of the wall.

[0103] A further embodiment of any of the preceding or following rocket engine system embodiments, comprising:

[0104] An embodiment further comprising at least one port positioned to apply the coolant fuel to a first surface of the wall.

[0105] A further embodiment of any of the preceding or following rocket engine system embodiments, wherein the oxidizer comprises liquid oxygen.

[0106] A further embodiment of any of the preceding or following rocket engine system embodiments, wherein the oxidizer comprises air.

[0107] Further embodiments of any of the preceding or following rocket engine system embodiments, wherein the oxidizer comprises hydrogen peroxide.

[0108] In a further embodiment of any of the preceding or following rocket engine system embodiments, the wall comprises:

[0109] further having at least one cooling passage formed therein;

[0110] An embodiment wherein the coolant source is fluidly connected to the at least cooling passage.

[0111] A further embodiment of the technology herein is a method for improving efficiency of a propulsion engine, comprising:

[0112] applying a coolant fuel along a first surface of a wall partially defining a combustion chamber of the propulsion engine, the coolant fuel film cooling the first surface of the wall;

[0113] applying an oxidizer to a second surface of the wall partially defining the combustion chamber;

[0114] monitoring the coolant fuel flow and the oxidant flow; and

[0115] controlling the flow of the coolant fuel and the flow of the oxidizer so that a stoichiometric ratio of the coolant fuel and the oxidizer mixture is suitable for causing combustion of the coolant fuel and the oxidizer mixture after the coolant fuel has film-cooled the first surface of the wall, thereby increasing thrust of the propulsion engine by combustion of the coolant fuel and the oxidizer mixture, wherein the increased thrust increases efficiency of the propulsion engine.

[0116] A further embodiment of any of the preceding or following method embodiments, wherein applying the coolant fuel along a first surface of the wall provides a protective barrier to the first surface of the wall.

[0117] In a further embodiment of any of the preceding or following method embodiments,

[0118] Embodiments further comprising applying the coolant fuel to a first surface of the wall using at least one port proximate to the first surface of the wall.

[0119] Further embodiments of any of the preceding or following method embodiments, wherein applying the oxidizer to the second surface of the wall comprises applying liquid oxygen to the second surface of the wall.

[0120] Further embodiments of any of the preceding or following method embodiments, wherein applying the oxidant to the second surface of the wall comprises applying air to the second surface of the wall.

[0121] Further embodiments of any of the preceding or following method embodiments, wherein applying the oxidizing agent to the second surface of the wall comprises applying hydrogen peroxide to the second surface of the wall.

[0122] In a further embodiment of any of the above-described rocket engine system embodiments, applying the coolant fuel to a first surface of the wall comprises:

[0123] An embodiment, further comprising applying the coolant fuel to a first surface of the wall using at least one cooling passage formed in the wall.

Claims

1. a coolant source including a coolant fuel configured to supply the coolant fuel to a first surface of a wall that partially defines an annulus in which detonation occurs in the rotary detonation engine, the coolant fuel film cooling the first surface of the wall; an oxidant source configured to supply oxidant to a second surface of the wall that partially defines the annulus; and a monitor configured to control the flow of the coolant fuel and the flow of the oxidizer, the monitor being further configured to ensure that a stoichiometric ratio of the mixture of the coolant fuel and the oxidizer is suitable for causing combustion of the mixture of the coolant fuel and the oxidizer after the coolant fuel has film cooled the first surface of the wall. A detonation engine system having:

2. 2. The rocket engine system of claim 1, wherein said coolant fuel film cools said first surface of said wall, thereby providing a protective barrier to said first surface of said wall.

3. 2. The rocket engine system of claim 1, further comprising at least one port positioned for applying said coolant fuel to said first surface of said wall.

4. 10. The rocket engine system of claim 1, wherein said oxidizer comprises liquid oxygen.

5. 10. The rocket engine system of claim 1, wherein said oxidizer comprises air.

6. 10. The rocket engine system of claim 1, wherein said oxidizer comprises hydrogen peroxide.

7. The wall portion is further having at least one cooling passage formed therein; The rocket engine system of claim 1 , wherein said coolant source is fluidly connected with said at least one cooling passage.

8. a coolant source including a coolant fuel, the coolant source configured to supply the coolant fuel to a first surface of a wall that partially defines a combustion chamber of the rocket engine, the coolant fuel providing film cooling to the first surface of the wall; an oxidizer source configured to supply oxidizer to a second surface of the wall that partially defines the combustion chamber; and a monitor configured to control the flow of the coolant fuel and the flow of the oxidizer, the monitor being further configured to ensure that a stoichiometric ratio of the mixture of the coolant fuel and the oxidizer is suitable for causing combustion of the mixture of the coolant fuel and the oxidizer after the coolant fuel has film cooled the first surface of the wall. A rocket engine system having:

9. 9. The rocket engine system of claim 8, wherein said coolant fuel film cools said first surface of said wall, thereby providing a protective barrier to said first surface of said wall.

10. 9. The rocket engine system of claim 8, further comprising at least one port positioned for applying said coolant fuel to said first surface of said wall.

11. 9. A rocket engine system as recited in claim 8, wherein said oxidizer comprises liquid oxygen.

12. 9. A rocket engine system as recited in claim 8, wherein said oxidizer comprises air.

13. 9. A rocket engine system as recited in claim 8, wherein said oxidizer comprises hydrogen peroxide.

14. The wall portion is further having at least one cooling passage formed therein; 9. The rocket engine system of claim 8, wherein said coolant source is fluidly connected with said at least one cooling passage.

15. 1. A method for improving efficiency of a propulsion engine, comprising: applying a coolant fuel along a first surface of a wall partially defining a combustion chamber of the propulsion engine, the coolant fuel film cooling the first surface of the wall; applying an oxidizer to a second surface of the wall partially defining the combustion chamber; monitoring the coolant fuel flow and the oxidant flow; and controlling the flow of the coolant fuel and the flow of the oxidizer so that a stoichiometric ratio of the coolant fuel and the oxidizer mixture is suitable for causing combustion of the coolant fuel and the oxidizer mixture after the coolant fuel has film-cooled the first surface of the wall, thereby increasing thrust of the propulsion engine by combustion of the coolant fuel and the oxidizer mixture, wherein the increased thrust increases efficiency of the propulsion engine. A method having the following.

16. The method of claim 15 , further comprising applying the coolant fuel along a first surface of the wall to provide a protective barrier to the first surface of the wall.

17. The method of claim 15 , further comprising applying the coolant fuel to the first surface of the wall using at least one port proximate to the first surface of the wall.

18. The method of claim 15 , wherein applying the oxidizer to the second surface of the wall comprises applying liquid oxygen to the second surface of the wall.

19. The method of claim 15 , wherein applying the oxidizer to the second surface of the wall comprises applying air to the second surface of the wall.

20. 16. The method of claim 15, wherein applying the oxidizing agent to the second surface of the wall comprises applying hydrogen peroxide to the second surface of the wall.

21. applying the coolant fuel to a first surface of the wall; 9. The rocket engine system of claim 8, further comprising applying said coolant fuel to a first surface of said wall using at least one cooling passage formed in said wall.

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