Patent Publication Number: US-2018038316-A1

Title: Liquid-fueled rocket engine assemblies, and related methods of using liquid-fueled rocket engine assemblies

Description:
TECHNICAL FIELD 
     Embodiments of the disclosure relate generally to liquid-fueled rocket engines assemblies, and to methods of using liquid-fueled rocket engine assemblies. More particularly, embodiments of the disclosure relate to liquid-fueled rocket engine assemblies including expander cycle configurations, and to methods of using such liquid-fueled rocket engine assemblies. 
     BACKGROUND 
     Liquid-fueled rocket engine assemblies utilize liquids as one or more of propellant sources, fuel sources, and oxidizer sources. Liquid-fueled rocket engine assemblies can be quickly fueled and refueled, and the relatively high density of liquids can facilitate the use of relatively smaller storage vessels. 
     One example, of a liquid-fueled rocket engine assembly is an expander cycle rocket engine assembly. An expander cycle rocket engine assembly heats liquid fuel to form a gaseous fuel that is utilized to drive a turbine to power one or more pumps of the expander cycle rocket engine assembly. Expander cycle rocket engine assemblies do not require a gas generator or pre-burner. Some expander cycle rocket engine assemblies have a closed-cycle configuration wherein all of the gaseous fuel exiting the turbine is directed into a combustion chamber of the rocket engine assembly. However, such a configuration generally requires a relatively high pump discharge to combustion chamber pressure ratio between 2.5 and 3 or higher, which necessitates relatively high turbine power and turbine inlet pressure and can result in higher assembly weight and lower reliability. Other expander cycle rocket engine assemblies have an open-cycle configuration wherein a portion of the liquid fuel is heated and used to drive a turbine and is then discharged (i.e., is not directed into a combustion chamber of the rocket engine assembly). Such a configuration can reduce cycle peak pressures as compared to the previously mentioned closed-cycle configuration, but discharging some of the liquid fuel results in efficiency losses. Further expander cycle rocket engine assemblies exhibit a closed-cycle configuration wherein all of the gaseous fuel exiting the turbine is cooled, condensed, and mixed with additional liquid fuel. Such a configuration exhibits the high performance of the previously mentioned closed-cycle configuration and the reduced cycle peak pressure of the previously mentioned open-cycle configuration, but requires additional equipment (e.g., heat exchangers, condensers, mixers, etc.) that can be prohibitively large and heavy to facilitate various operational parameters (e.g., higher combustion chamber pressures, higher fluid flow rates, higher turbine power outputs, etc.) desirable or necessary for various applications of the expander rocket engine assemblies. In addition, such a configuration may be unable to support relatively higher chamber pressures when recirculation flow is unable to the fully condensed. 
     It would, therefore, be desirable to have new liquid-fueled rocket engine assemblies and related methods that alleviate one or more of the above problems. 
     BRIEF SUMMARY 
     Embodiments described herein include liquid-fueled rocket engines assemblies, and methods of using the liquid-fueled rocket engine assemblies. In some embodiments, a liquid-fueled rocket engine assembly comprises a combustor assembly, a combustor jacket, a turbine, a flow control device, a discharge device, a heat exchanger, and a mixer. The combustor assembly comprises an injector, a combustion chamber, and a nozzle. The combustor jacket at least partially surrounds and is configured to cool the combustion chamber and the primary nozzle of the combustor assembly using a portion of a pressurized liquid fuel exiting a fuel pump. The turbine is configured and positioned to receive and expand a gaseous fuel exiting the combustor jacket to power one or more of the fuel pump and an oxidizer pump. The flow control device comprises a first outlet and a second outlet, and is configured and positioned to direct an expanded gaseous fuel exiting the turbine through one or more of the first outlet and the second outlet. The discharge device is configured and positioned to receive, expand, and exhaust at least a portion of the expanded gaseous fuel exiting the first outlet of the flow control device. The heat exchanger is configured and positioned to heat another portion of the pressurized liquid fuel exiting the fuel pump with at least a portion of the expanded gaseous fuel exiting the second outlet of the flow control device and to direct a resulting heated, pressurized liquid fuel into the combustor assembly. The mixer is configured and positioned to combine a cooled fuel exiting the heat exchanger with a liquid fuel exiting a fuel source to form a mixed liquid fuel and to direct the mixed liquid fuel to the fuel pump. 
     In additional embodiments, a liquid-fueled rocket engine assembly comprises an oxidizer pump, a fuel pump, a coolant pump, a combustor assembly, a combustor jacket, a turbine, and a discharge device. The oxidizer pump is configured and positioned to pressurize a liquid oxidizer exiting a coolant source. The fuel pump is configured and positioned to pressurize a liquid fuel exiting a fuel source. The coolant pump is configured and positioned to pressurize a liquid coolant exiting a coolant source. The combustor assembly comprises an injector configured and positioned to receive and combine the pressurized liquid oxidizer and the pressurized liquid fuel to form a reactant mixture, a combustion chamber configured and positioned to receive and combust the reactant mixture to produce propellant gases, and a primary nozzle configured and positioned to receive and exhaust the propellant gases. The combustor jacket surrounds and is configured to cool at least a portion of the combustion chamber and the primary nozzle of the combustor assembly using the pressurized liquid coolant exiting the coolant pump. The turbine is configured and positioned to receive and expand a gaseous coolant exiting the combustor jacket to power one or more of the coolant pump, the fuel pump, and the oxidizer pump. The discharge device is configured and positioned to receive, expand, and exhaust the expanded gaseous coolant exiting the turbine. 
     In yet further embodiments, a method of using a liquid-fueled rocket engine assembly comprises directing a first pressurized liquid material and a second pressurized liquid material into a combustor assembly comprising an injector, a combustion chamber, and a primary nozzle. A third pressurized liquid material is directed into a combustor jacket surrounding at least a portion of the combustor assembly to cool the at least a portion of the combustor assembly and form a gaseous material. The gaseous material is expanded within a turbine operatively associated with at least one pump used to form one or more of the first pressurized liquid material, the second pressurized liquid material, and the third pressurized liquid material to drive the turbine and power the at least one pump. From about 0 percent to about 100 percent of an expanded gaseous material exiting the turbine is exhausted. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  shows a simplified schematic representation of a liquid-fueled rocket engine assembly, according to an embodiment of the disclosure. 
         FIG. 2  shows a simplified schematic representation of a liquid-fueled rocket engine assembly, according to another embodiment of the disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     Liquid-fueled rocket engines assemblies are described, as are methods of using the liquid-fueled rocket engine assemblies. For example, in accordance with an embodiment of the disclosure, a liquid-fueled rocket engine assembly includes a combustor assembly, an oxidizer source, an oxidizer pump, a fuel source, a fuel pump, a combustor jacket, a turbine, a flow control device, a discharge device, a heat exchanger, and a mixer. The combustor assembly includes an injector, a combustion chamber, and a nozzle. The oxidizer pump is configured and positioned to pressurize a liquid oxidizer (e.g., liquid oxygen, liquid hydrogen peroxide, etc.) exiting the oxidizer source and to direct the pressurized liquid oxidizer into the combustor assembly. The combustor jacket surrounds the combustor assembly and is configured to cool the combustor assembly using a portion of a pressurized liquid fuel exiting the fuel pump. The turbine is configured and positioned to receive and expand a gaseous fuel exiting the combustor jacket to power one or more (e.g., each) of the fuel pump and the oxidizer pump. The flow control device comprises a first outlet and a second outlet, and is configured and positioned to direct an expanded gaseous fuel exiting the turbine through one or more of the first outlet and the second outlet. The flow control device may adjustably control amounts of the expanded gaseous fuel exiting the first outlet and the second outlet. The amounts of the expanded gaseous fuel exiting the first outlet and the second outlet may be selected at least partially based on properties (e.g., material composition, flow rate, temperature, pressure, etc.) of the expanded gaseous fuel and on desired operational parameters (e.g., desired turbine power output, desired combustion chamber pressure, etc.) of the liquid-fueled rocket engine assembly to maintain an expander cycle of the liquid-fueled rocket engine assembly. The discharge device is configured and positioned to receive, expand, and exhaust at least a portion of the expanded gaseous fuel (if any) exiting the first outlet of the flow control device. The heat exchanger is configured and positioned to heat another portion of the pressurized liquid fuel exiting the fuel pump with at least a portion of the expanded gaseous fuel (if any) exiting the second outlet of the flow control device and to direct a resulting heated, pressurized liquid fuel into the combustor assembly. The mixer is configured and positioned to combine a cooled fuel exiting the heat exchanger with a liquid fuel (e.g., liquid hydrogen; liquid ammonia; a low molecular weight liquid hydrocarbon, such as liquid methane; etc.) exiting the fuel source to form a mixed liquid fuel and to direct the mixed liquid fuel to the fuel pump. The liquid-fueled rocket engine assemblies and methods of the disclosure may reliably establish and maintain expander cycles for a wide variety of materials (e.g., fuels, oxidizers, coolants, etc.) and process parameters (e.g., flow rates, combustion chamber pressures, temperatures, etc.). The liquid-fueled rocket assemblies and methods of the disclosure may provide increased flexibility, increased efficiency, and/or reduced costs as compared to conventional liquid-fueled rocket engine assemblies and conventional methods of using liquid-fueled rocket engine assemblies. 
     The following description provides specific details, such as sizes, shapes, material compositions, and orientations in order to provide a thorough description of embodiments of the disclosure. However, a person of ordinary skill in the art would understand that the embodiments of the disclosure may be practiced without necessarily employing these specific details. Embodiments of the disclosure may be practiced in conjunction with conventional fabrication techniques employed in the industry. In addition, the description provided below does not foini a complete process flow for manufacturing a liquid-fueled rocket engine assembly. Only those process acts and structures necessary to understand the embodiments of the disclosure are described in detail below. Additional acts to form a complete liquid-fueled rocket engine assembly from the structures described herein may be performed by conventional fabrication processes. 
     Drawings presented herein are for illustrative purposes only, and are not meant to be actual views of any particular material, component, structure, device, or system. Variations from the shapes depicted in the drawings as a result, for example, of manufacturing techniques and/or tolerances, are to be expected. Thus, embodiments described herein are not to be construed as being limited to the particular shapes or regions as illustrated, but include deviations in shapes that result, for example, from manufacturing. For example, a region illustrated or described as box-shaped may have rough and/or nonlinear features, and a region illustrated or described as round may include some rough and/or linear features. Moreover, sharp angles that are illustrated may be rounded, and vice versa. Thus, the regions illustrated in the figures are schematic in nature, and their shapes are not intended to illustrate the precise shape of a region and do not limit the scope of the present claims. The drawings are not necessarily to scale. 
     As used herein, the terms “comprising,” “including,” “containing,” “characterized by,” and grammatical equivalents thereof are inclusive or open-ended terms that do not exclude additional, unrecited elements or method acts, but also include the more restrictive terms “consisting of” and “consisting essentially of” and grammatical equivalents thereof. As used herein, the term “may” with respect to a material, structure, feature or method act indicates that such is contemplated for use in implementation of an embodiment of the disclosure and such term is used in preference to the more restrictive term “is” so as to avoid any implication that other, compatible materials, structures, features and methods usable in combination therewith should or must be, excluded. 
     As used herein, spatially relative terms, such as “beneath,” “below,” “lower,” “bottom,” “above,” “over,” “upper,” “top,” “front,” “rear,” “left,” “right,” and the like, may be used for ease of description to describe one element&#39;s or feature&#39;s relationship to another element(s) or feature(s) as illustrated in the figures. Unless otherwise specified, the spatially relative terms are intended to encompass different orientations of the materials in addition to the orientation depicted in the figures. For example, if materials in the figures are inverted, elements described as “over” or “above” or “on” or “on top of” other elements or features would then be oriented “below” or “beneath” or “under” or “on bottom of” the other elements or features. Thus, the term “over” can encompass both an orientation of above and below, depending on the context in which the term is used, which will be evident to one of ordinary skill in the art. The materials may be otherwise oriented (e.g., rotated 90 degrees, inverted, flipped) and the spatially relative descriptors used herein interpreted accordingly. 
     As used herein, the singular forms “a,” “an,” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. 
     As used herein, “and/or” includes any and all combinations of one or more of the associated listed items. 
     As used herein, the terms “configured” and “configuration” refer to a size, shape, material composition, orientation, and arrangement of one or more of at least one structure and at least one apparatus facilitating operation of one or more of the structure and the apparatus in a predetermined way. 
     As used herein, the term “substantially” in reference to a given parameter, property, or condition means and includes to a degree that one of ordinary skill in the art would understand that the given parameter, property, or condition is met with a degree of variance, such as within acceptable manufacturing tolerances. By way of example, depending on the particular parameter, property, or condition that is substantially met, the parameter, property, or condition may be at least 90.0% met, at least 95.0% met, at least 99.0% met, or even at least 99.9% met. 
     As used herein, the term “about” in reference to a given parameter is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the given parameter). 
       FIG. 1  is a simplified schematic view of a liquid-fueled rocket engine assembly  100 , in accordance with an embodiment of the disclosure. As shown in  FIG. 1 , the liquid-fueled rocket engine assembly  100  includes a combustor assembly  102  including an injector  104 , a combustion chamber  106 , and a nozzle  108 . The liquid-fueled rocket engine assembly  100  also includes at least one oxidizer source  110 , at least one oxidizer pump  114  downstream of the oxidizer source  110  and upstream of the combustor assembly  102 , at least one fuel source  112 , at least one fuel pump  116  downstream of the fuel source  112 , at least one turbine  118  operatively associated with (e.g., mechanically coupled to) the oxidizer pump  114  and the fuel pump  116 , at least one flow control device  142  downstream of the turbine  118 , at least one discharge device  120  downstream of the flow control device  142 , at least one heat exchanger  122  downstream of each of the flow control device  142  and the fuel pump  116  and upstream of the combustor assembly  102 , at least one mixer  124  downstream of each of the heat exchanger  122  and the fuel source  112  and upstream of the fuel pump  116 , at least one combustor jacket  126  downstream of the mixer  124  and upstream of the turbine  118 , and various conduits configured and positioned to direct one or more fluids between the various components of the liquid-fueled rocket engine assembly  100 . 
     The oxidizer source  110  may comprise at least one vessel (e.g., at least one pressure vessel) configured and operated to at least temporarily hold (e.g., store, contain, etc.) a volume of liquid oxidizer, such as liquid oxygen, liquid hydrogen peroxide, etc. The oxidizer source  110  may be fluidly coupled to an inlet of the oxidizer pump  114 , such that oxidizer exiting the oxidizer source  110  is directed (e.g., flowed, fed, delivered, etc.) into the oxidizer pump  114 . In addition, at least one oxidizer boost pump may, optionally, be positioned between the oxidizer source  110  and the oxidizer pump  114 . If present, the oxidizer boost pump may be operatively associated with at least one oxidizer boost turbine, and may be configured to pump liquid oxidizer from the oxidizer source  110  into the oxidizer pump  114 . 
     The oxidizer pump  114  may comprise at least one device or apparatus configured and positioned to receive liquid oxidizer (e.g., liquid oxygen, liquid hydrogen peroxide, etc.) from the oxidizer source  110  and to increase the pressure of the liquid oxidizer. An outlet of the oxidizer pump  114  may be fluidly coupled to (e.g., in fluid communication with) an inlet of the injector  104  of the combustor assembly  102 , such that pressurized liquid oxidizer exiting the oxidizer pump  114  is directed into the injector  104 , as described in further detail below. Optionally, as shown in  FIG. 1 , an oxidizer control valve  128  may be positioned between the oxidizer pump  114  and the injector  104  of the combustor assembly  102 , and may be configured to control an amount (e.g., mass flow) of liquid oxidizer directed into the injector  104 . 
     The fuel source  112  may comprise at least one vessel (e.g., at least one pressure vessel) configured and operated to at least temporarily hold (e.g., store, contain, etc.) a volume of liquid fuel, such as liquid hydrogen, liquid ammonia, a low molecular weight liquid hydrocarbon (e.g., liquid methane), combinations thereof, etc. The fuel source  112  may be fluidly coupled to an inlet of the mixer  124 , such that liquid fuel exiting the fuel source  112  is directed into the mixer  124 . In addition, as shown in  FIG. 1 , at least one fuel boost pump  130  may, optionally, be positioned between the fuel source  112  and the mixer  124 . If present, the fuel boost pump  130  may be operatively associated with at least one fuel boost turbine  132 , and may be configured to pump liquid fuel from fuel source  112  into the mixer  124 . 
     The mixer  124  may comprise at least one device or apparatus configured and positioned to receive each of liquid fuel (e.g., liquid hydrogen, liquid ammonia, liquid methane, etc.) from the fuel source  112  and additional (e.g., recirculated) fuel from the heat exchanger  122  and to produce a mixed liquid fuel that may be directed to the fuel pump  116 . The liquid fuel received from the fuel source  112  may be relatively cold as compared to the additional fuel received from the heat exchanger  122 . Accordingly, the mixer  124  may be configured to sufficiently mix the liquid fuel received from the fuel source  112  and the additional fuel received from the heat exchanger  122  to prevent the formation of gas pockets within the mixer  124  and also prevent cavitation within the fuel pump  116 . 
     The fuel pump  116  may comprise at least one device or apparatus configured and positioned to receive liquid fuel (e.g., mixed liquid fuel) from the mixer  124  and to increase the pressure of the fuel. As shown in  FIG. 1 , the fuel pump  116  may include outlets fluidly coupled to (e.g., in fluid communication with) inlets of the combustor jacket  126  and the heat exchanger  122 , such that pressurized liquid fuel exiting the fuel pump  116  is directed into one or more of the combustor jacket  126  and a first channel  134  of the heat exchanger  122 , as described in further detail below. In addition, as shown in  FIG. 1 , in some embodiments wherein the liquid-fueled rocket engine assembly  100  includes a fuel boost pump  130  and an associated fuel boost turbine  132 , a control valve  138  may, optionally, be positioned between the fuel pump  116  and each of the combustor jacket  126  and the fuel boost turbine  132 . The control valve  138  may be configured to control (e.g., increase or decrease) amounts of the pressurized liquid fuel directed into the combustor jacket  126  and the fuel boost turbine  132 . 
     The combustor jacket  126  may comprise at least one device or apparatus configured and positioned to receive pressurized liquid fuel from the fuel pump  116  and to transfer heat from components of the combustor assembly  102  to the pressurized liquid fuel. As shown in  FIG. 1 , the combustor jacket  126  may at least partially (e.g., substantially) surround the combustion chamber  106  of the combustor assembly  102  and the nozzle  108  of the combustor assembly  102 . In addition, an outlet of the combustor jacket  126  may be fluidly coupled to (e.g., in fluid communication with) an inlet of the turbine  118 . As heat is transferred from the combustor assembly  102  to the pressurized liquid fuel within the combustor jacket  126 , the pressurized liquid fuel may transition from a liquid phase to a gaseous phase (e.g., a vapor phase), and then the heated, gaseous fuel may be directed to the turbine  118 . The combustor jacket  126  may be configured to extract sufficient heat from the combustor assembly  102  to prevent overheating and damage to the components of the combustor assembly  102 . 
     The turbine  118  may comprise at least one device or apparatus configured and positioned to drive the oxidizer pump  114  and the fuel pump  116 . The turbine  118  may be configured and positioned to receive heated, gaseous fuel from the combustor jacket  126 , and to convert thermal energy stored in the heated, gaseous fuel into kinetic energy to drive one or more (e.g., each) of the oxidizer pump  114  and the fuel pump  116 . The turbine  118  may be mechanically coupled to one or more (e.g., each) of the oxidizer pump  114  and the fuel pump  116 . For example, as shown in  FIG. 1 , the turbine  118 , the oxidizer pump  114 , and the fuel pump  116  may share a single (e.g., only one) shaft  140  that directly, rotationally couples the turbine  118 , the oxidizer pump  114 , and the fuel pump  116 . In some embodiments, a transmission (e.g., a gear box) may be used to mechanically couple the turbine  118  to the shaft  140  and transfer power from the turbine  118  to the oxidizer pump  114  and/or the fuel pump  116 . In further embodiments, the oxidizer pump  114  and the fuel pump  116  may be operatively associated with (e.g., mechanically coupled to) separate turbines. An outlet of the turbine  118  may be fluidly coupled to (e.g., in fluid communication with) an inlet of a flow control device  142  (e.g., three-way valve), such that a cooled, expanded gaseous fuel exiting the turbine  118  is directed into the flow control device  142 . 
     The flow control device  142  may comprise at least one device or apparatus configured and positioned to receive the cooled, expanded gaseous fuel exiting the turbine  118 , and to control amounts (e.g., mass flow) of the cooled, expanded gaseous fuel directed to each of the discharge device  120  and the heat exchanger  122 . The flow control device  142  may include outlets fluidly coupled to (e.g., in fluid communication with) inlets of the discharge device  120  and the heat exchanger  122 , as well as means of controlling fluid flow through each of the outlets, such that the cooled, expanded gaseous fuel received by the flow control device  142  is controllably directed into one or more of the discharge device  120  and a second channel  136  of the heat exchanger  122 . By way of non-limiting example, the flow control device  142  may comprise a three-way valve. A ratio of an amount of the cooled, expanded gaseous fuel directed to the discharge device  120  to another amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  may be selected at least partially based on properties (e.g., material composition, flow rate, temperature, pressure, etc.) of the cooled, expanded gaseous fuel and on desired operational parameters (e.g., a desired power output of the turbine  118 , a desired chamber pressure of the combustion chamber  106 , etc.) and predetermined performance metrics of the liquid-fueled rocket engine assembly  100 . The flow control device  142  may, for example, be configured and positioned to direct (e.g., divert) from about 0 percent to about 100 percent (e.g., from about 0 percent to about 50 percent, such as 0 percent, or from about 20 percent to about 50 percent; from about or from about 50 percent to about 100 percent, such as 100 percent, or from about 50 percent to about 80 percent) of the cooled, expanded gaseous fuel exiting the turbine  118  to the heat exchanger  122 , and to direct a remainder of the cooled, expanded gaseous fuel exiting the turbine  118  to the discharge device  120 . 
     The discharge device  120  may comprise at least one device or apparatus configured and positioned to receive, expand, and discharge (e.g., exhaust) at least a portion of the cooled, expanded gaseous fuel exiting the flow control device  142  (e.g., a portion of the cooled, expanded gaseous fuel not directed to the heat exchanger  122 ). As shown in  FIG. 1 , in some embodiments, the discharge device  120  comprises a secondary nozzle separate (e.g., discrete) from the combustor assembly  102 . In additional embodiments, the discharge device  120  comprises a component of the combustor assembly  102  configured and operated to receive and discharge the cooled, expanded gaseous fuel. For example, the discharge device  120  may comprise a portion of the nozzle  108 , such as a portion of the nozzle  108  configured and operated to facilitate supersonic flow therein. In such embodiments, the cooled, expanded gaseous fuel may be directed into the component (e.g., the nozzle  108 ) of the combustor assembly  102  without passing through and/or being acted upon (e.g., combusted) by one or more other components (e.g., the combustion chamber  106 ) of the combustor assembly  102 . The ability to discharge at least a portion of the cooled, expanded gaseous fuel exiting the flow control device  142  may permit the liquid-fueled rocket engine assembly  100  to achieve desirable operational parameters (e.g., relatively higher turbine power outputs, relatively higher combustion chamber pressures, etc.) for a variety of different fuel types, oxidizer types, fuel flowrates, and/or oxidizer flowrates without performance losses and/or failures that may otherwise be associated with configuring (e.g., sizing) one or more components (e.g., heat exchangers, mixers, etc.) of the liquid-fueled rocket engine assembly  100  to accommodate (e.g., receive) all of the cooled, expanded gaseous fuel exiting the turbine  118 . 
     The heat exchanger  122  may comprise at least one device or apparatus configured and positioned to receive a portion of the pressurized liquid fuel exiting the fuel pump  116  (e.g., a portion of the pressurized liquid fuel exiting the fuel pump  116  not directed to at least one of the combustor jacket  126  and the fuel boost turbine  132 ) and at least a portion of the cooled, expanded gaseous fuel exiting the flow control device  142  (e.g., a portion of the cooled, expanded gaseous fuel not directed to the discharge device  120 ), and to facilitate the transfer of heat from the cooled, expanded gaseous fuel from the flow control device  142  to the pressurized liquid fuel from the fuel pump  116 . By way of non-limiting example, the heat exchanger  122  may be configured as one or more of a counter-flow heat exchanger, a shell and tube heat exchanger, a plate heat exchanger, and a plate fin heat exchanger. As shown in  FIG. 1 , the portion of the pressurized liquid fuel may be received into the first channel  134  of the heat exchanger  122 , and the at least a portion of the cooled, expanded gaseous fuel may be received into the second channel  136  of the heat exchanger  122  fluidly separated (e.g., fluidly isolated) from the first channel  134 . Within the second channel  136  of the heat exchanger  122 , the cooled, expanded gaseous fuel received from the flow control device  142  may be further cooled to a temperature close to a phase-transition temperature thereof to form additional fuel, which may exit the second channel  136  of the heat exchanger  122  and be directed into the mixer  124  to be combined (e.g., mixed) with the liquid fuel from the fuel source  112  to form additional mixed liquid fuel. Optionally, another control valve  144  may be positioned between the heat exchanger  122  and the mixer  124 , and may be configured to control an amount of the additional fuel directed into the mixer  124 . Furthermore, within the first channel  134  of the heat exchanger  122 , the portion of the pressurized liquid fuel from the fuel pump  116  may be heated to form a heated, pressurized liquid fuel, which may exit the first channel  134  of the heat exchanger  122  and be directed into the injector  104  of the combustor assembly  102 . 
     The injector  104  of the combustor assembly  102  may comprise at least one device or apparatus configured and positioned to mix the heated, pressurized liquid fuel received from the heat exchanger  122  and the pressurized liquid oxidizer received from the oxidizer pump  114 , and to deliver the resulting reactant mixture to the combustion chamber  106  of the combustor assembly  102 , wherein the reactant mixture may be ignited and combusted to produce propellant gases. In turn, the nozzle  108  may be coupled to the combustion chamber  106  and may be configured to direct the produced propellant gases out of the combustor assembly  102  through an opening at an end of the nozzle  108  to generate thrust (e.g., sufficient thrust to propel a vehicle including the liquid-fueled rocket engine assembly  100 ). 
     During use and operation of the liquid-fueled rocket engine assembly  100 , a liquid fuel (e.g., liquid hydrogen, liquid ammonia, liquid methane, etc.) may be directed from the fuel source  112  into the mixer  124 . Within the mixer  124 , the liquid fuel may, optionally, be combined with additional (e.g., recirculated) fuel received from the heat exchanger  122 . The liquid fuel and/or resulting mixed liquid fuel may then exit the mixer  124  and may be delivered into the fuel pump  116 . Thereafter, a resulting pressurized liquid fuel may be separated (e.g., split, divided, etc.) into a first portion (e.g., a first fuel substream) and a second portion (e.g., a second fuel substream). The first portion may be directed through the first channel  134  of the heat exchanger  122  and into the injector  104  of the combustor assembly  102  using pressure provided by the fuel pump  116 . Within the injector  104 , the first portion of the pressurized liquid fuel may be combined with pressurized liquid oxidizer delivered from the oxidizer source  110  using the oxidizer pump  114  to produce a reactant mixture. The reactant mixture may be directed into and combusted within the combustion chamber  106  of the combustor assembly  102  to generate propellant gases, which may then be exhausted from the nozzle  108  of the combustor assembly  102  to generate thrust. Concurrently, the second portion of the pressurized liquid fuel may be directed into the combustor jacket  126 . Within the combustor jacket  126 , the second portion of the pressurized liquid fuel may extract heat from the combustor assembly  102  to form a heated, gaseous fuel that may be directed into the turbine  118  to drive the turbine  118  and power one or more (e.g., each) of the oxidizer pump  114  and the fuel pump  116 . Cooled, expanded gaseous fuel may then exit the turbine  118  and may be directed into the flow control device  142 . Within the flow control device  142 , the cooled, expanded gaseous fuel may, optionally, be divided into a first portion (e.g., a first cooled, expanded gaseous fuel substream) and a second portion (e.g., a second cooled, expanded gaseous fuel substream) at least partially depending on properties (e.g., material composition, flow rate, temperature, pressure, etc.) of the cooled, expanded gaseous fuel and on desired operational parameters (e.g., turbine power output, combustion chamber pressure, etc.) of the liquid-fueled rocket engine assembly  100 , as described in further detail below. The first portion (if any) of the cooled, expanded gaseous fuel may be directed into the discharge device  120 , wherein it may be expanded and exhausted. The second portion (if any) of the cooled, expanded gaseous fuel may be directed through the second channel  136  of the heat exchanger  122  and into the mixer  124  and may be combined with more fuel from the fuel source  112 . 
     The relative amounts of the cooled, expanded gaseous fuel directed to the discharge device  120  and the heat exchanger  122  may be controlled (e.g., selected, maintained, adjusted, etc.) to ensure that the heat exchanger  122  and/or the mixer  124  is/are able to sufficiently decrease the temperature of and condense the cooled, expanded gaseous fuel to maintain an expander cycle of the liquid-fueled rocket engine assembly  100 . For example, in some embodiments where it is desirable to operate the liquid-fueled rocket engine assembly  100  at a relatively higher combustion chamber pressure (e.g., a combustion chamber pressure greater than or equal to about 1200 psi, or a combustion chamber pressure greater than or equal to about 1600 psi), the flow control device  142  may be used to decrease an amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  (and, hence, increase an amount of the cooled, expanded gaseous fuel directed into the discharge device  120 ). Without decreasing the amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  (e.g., such as if all of the cooled, expanded gaseous fuel entering the flow control device  142  was directed to the heat exchanger  122 ), the liquid-fueled rocket engine assembly  100  may be unable to maintain the expander cycle at the material flow rates and associated pump power requirements needed to facilitate the relatively higher combustion chamber pressure or may require an undesirably large heat exchanger  122  (and/or mixer  124 ) to facilitate sufficient additional cooling of the cooled, expanded gaseous fuel. By way of non-limiting example, an amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  from the flow control device  142  may be within a range of from about 0 percent to about 50 percent of an amount of the cooled, expanded gaseous fuel entering into the flow control device  142 . In some embodiments, an amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  from the flow control device  142  is within a range of from about 20 percent to about 50 percent of the amount of the cooled, expanded gaseous fuel entering into the flow control device  142 . In further embodiments, an amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  from the flow control device  142  is about 0 percent of the amount of the cooled, expanded gaseous fuel entering into the flow control device  142 . 
     In additional embodiments, such as in embodiments where it is desirable to operate the liquid-fueled rocket engine assembly  100  at a relatively lower combustion chamber pressure (e.g., a combustion chamber pressure less than about 1200 psi), the flow control device  142  may be used to increase the amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  (and, hence, decrease the amount of the cooled, expanded gaseous fuel directed into the discharge device  120 ) to increase the efficiency and performance (e.g., by reducing the amount of exhausted fuel and improving specific impulse) of the liquid-fueled rocket engine assembly  100 . By way of non-limiting example, an amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  from the flow control device  142  may be within a range of from about 50 percent to about 100 percent of an amount of the cooled, expanded gaseous fuel entering into the flow control device  142 . In some embodiments, an amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  from the flow control device  142  is within a range of from about 50 percent to about 80 percent of the amount of the cooled, expanded gaseous fuel entering into the flow control device  142 . In further embodiments, an amount of the cooled, expanded gaseous fuel directed to the heat exchanger  122  from the flow control device  142  is about 100 percent of the amount of the cooled, expanded gaseous fuel entering into the flow control device  142 . 
       FIG. 2  is a simplified schematic view of a liquid-fueled rocket engine assembly  200  in accordance with an additional embodiment of the disclosure. As shown in  FIG. 2 , the liquid-fueled rocket engine assembly  200  includes a combustor assembly  202  including an injector  204 , a combustion chamber  206 , and a nozzle  208 . The liquid-fueled rocket engine assembly  200  also includes at least one oxidizer source  210 , at least one oxidizer pump  214  downstream of the oxidizer source  210  and upstream of the combustor assembly  202 , at least one fuel source  212 , at least one fuel pump  216  downstream of the fuel source  212  and upstream of the combustor assembly  202 , at least one coolant source  213 , at least one coolant pump  217  downstream of the coolant source  213 , at least one turbine  218  operatively associated with (e.g., mechanically coupled to) the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217 , at least one discharge device  220  downstream of the turbine  218 , at least one combustor jacket  226  downstream of the coolant pump  217  and upstream of the turbine  218 , and various conduits configured and positioned to direct one or more fluids between the various components of the liquid-fueled rocket engine assembly  200 . 
     The oxidizer source  210  may comprise at least one vessel (e.g., pressure vessel) configured and operated to at least temporarily hold (e.g., store, contain, etc.) a volume of liquid oxidizer, such as liquid oxygen, liquid hydrogen peroxide, combinations thereof, etc. The oxidizer source  210  may be fluidly coupled to an inlet of the oxidizer pump  214 , such that liquid oxidizer exiting the oxidizer source  210  is directed into the oxidizer pump  214 . In addition, at least one oxidizer boost pump may, optionally, be positioned between the oxidizer source  210  and the oxidizer pump  214 . If present, the oxidizer boost pump may be operatively associated with at least one oxidizer boost pump turbine, and may be configured to pump liquid oxidizer from the oxidizer source  210  into the oxidizer pump  214 . 
     The oxidizer pump  214  may comprise at least one device or apparatus configured and positioned to receive liquid oxidizer (e.g., liquid oxygen, liquid hydrogen peroxide, combinations thereof, etc.) from the oxidizer source  210  and to increase the pressure of the liquid oxidizer. An outlet of the oxidizer pump  214  may be fluidly coupled to (e.g., in fluid communication with) an inlet of the injector  204  of the combustor assembly  202 , such that pressurized liquid oxidizer exiting the oxidizer pump  214  is directed into the injector  204 . Optionally, as shown in  FIG. 2 , a control valve  228  may be positioned between the oxidizer pump  214  and the injector  204  of the combustor assembly  202 , and may be configured to control an amount (e.g., volume) of pressurized liquid oxidizer directed into the injector  204 . 
     The fuel source  212  may comprise at least one vessel (e.g., at least one pressure vessel) configured and operated to at least temporarily hold (e.g., store, contain, etc.) a volume of liquid fuel, such as liquid hydrogen, liquid ammonia, a liquid hydrocarbon (e.g., liquid methane, liquid propane, liquid kerosene, refined propellant-1 (RP-1), etc.), etc. The fuel source  212  may be fluidly coupled to an inlet of the fuel pump  216 , such that liquid fuel exiting the fuel source  212  is directed into the fuel pump  216 . In addition, at least one fuel boost pump may, optionally, be positioned between the fuel source  212  and the fuel pump  216 . If present, the fuel boost pump may be operatively associated with at least one fuel boost pump turbine, and may be configured to pump liquid fuel from fuel source  212  into the fuel pump  216 . 
     The fuel pump  216  may comprise at least one device or apparatus configured and positioned to receive liquid fuel (e.g., liquid hydrogen, liquid ammonia, a liquid hydrocarbon, etc.) from the fuel source  212  and to increase the pressure of the liquid fuel. An outlet of the fuel pump  216  may be fluidly coupled to (e.g., in fluid communication with) an inlet of the injector  204  of the combustor assembly  202 , such that pressurized liquid fuel exiting the fuel pump  216  is directed into the injector  204 . Optionally, as shown in  FIG. 2 , another control valve  244  may be positioned between the fuel pump  216  and the injector  204  of the combustor assembly  202 , and may be configured to control an amount (e.g., volume) of pressurized liquid fuel directed into the injector  204 . 
     The injector  204  of the combustor assembly  202  may comprise at least one device or apparatus configured and positioned to mix the pressurized liquid fuel received from the fuel pump  216  and the pressurized liquid oxidizer received from the oxidizer pump  214 , and to deliver the resulting reactant mixture to the combustion chamber  206  of the combustor assembly  202 , wherein the reactant mixture may be ignited and combust to produce propellant gases. In turn, the nozzle  208  may be coupled to the combustion chamber  206  and may be configured to direct the produced propellant gases out of the combustor assembly  202  through an opening at an end of the nozzle  208  to generate thrust (e.g., sufficient thrust to propel a vehicle including the liquid-fueled rocket engine assembly  200 ). 
     The coolant source  213  may comprise at least one vessel (e.g., at least one pressure vessel) configured and operated to at least temporarily hold (e.g., store, contain, etc.) a volume of liquid coolant, such as liquid hydrogen, liquid ammonia, a low molecular weight liquid hydrocarbon (e.g., liquid methane), liquid water, combinations thereof, etc. The coolant source  213  may be fluidly coupled to an inlet of the coolant pump  217 , such that liquid coolant exiting the coolant source  213  is directed into the coolant pump  217 . In addition, at least one coolant boost pump may, optionally, be positioned between the coolant source  213  and the coolant pump  217 . If present, the coolant boost pump may be operatively associated with at least one coolant boost turbine, and may be configured to pump liquid coolant from coolant source  213  into the coolant pump  217 . 
     The coolant pump  217  may comprise at least one device or apparatus configured and positioned to receive liquid coolant (e.g., liquid hydrogen, liquid ammonia, liquid methane, liquid water, etc.) from the coolant source  213  and to increase the pressure of the liquid coolant. An outlet of the coolant pump  217  may be fluidly coupled to (e.g., in fluid communication with) an inlet of the combustor jacket  226 , such that pressurized liquid coolant exiting the coolant pump  217  is directed into the combustor jacket  226 . Optionally, an additional control valve  246  may be positioned between the coolant pump  217  and the combustor jacket  226 , and may be configured to control an amount (e.g., volume) of pressurized liquid coolant directed into the combustor jacket  226 . 
     The combustor jacket  226  may comprise at least one device or apparatus configured and positioned to receive pressurized liquid coolant from the coolant pump  217  and to transfer heat from components of the combustor assembly  202  to the pressurized liquid coolant. As shown in  FIG. 2 , the combustor jacket  226  may at least partially (e.g., substantially) surround the combustion chamber  206  of the combustor assembly  202  and the nozzle  208  of the combustor assembly  202 . In addition, an outlet of the combustor jacket  226  may be fluidly coupled to (e.g., in fluid communication with) an inlet of the turbine  218 . As heat is transferred from the combustor assembly  202  to the pressurized liquid coolant within the combustor jacket  226 , the pressurized liquid coolant may transition from a liquid phase to a gaseous phase (e.g., a vapor phase), and then the heated, gaseous coolant may be directed into the turbine  218 . The combustor jacket  226  may be configured to extract sufficient heat from the combustor assembly  202  to prevent overheating and damage to the components of the combustor assembly  202 . 
     The turbine  218  may comprise at least one device or apparatus configured and positioned to drive the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217 . The turbine  218  may be configured and positioned to receive heated, gaseous coolant from the combustor jacket  226 , and to convert thermal energy stored in the heated, gaseous coolant into rotational movement of one or more (e.g., each) of the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217 . The turbine  218  may be mechanically coupled to one or more (e.g., each) of the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217 . For example, as shown in  FIG. 2 , the turbine  218 , the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217  may share a single (e.g., only one) shaft  240  that directly, rotationally couples the turbine  218 , the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217 . In some embodiments, a transmission (e.g., a gear box) may be used to mechanically couple the turbine  218  to the shaft  240  and transfer power from the turbine  218  to the oxidizer pump  214 , the fuel pump  216 , and/or the coolant pump  217 . In further embodiments, two or more of the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217  may be operatively associated with (e.g., mechanically coupled to) separate turbines. An outlet of the turbine  218  may be fluidly coupled to (e.g., in fluid communication with) an inlet of the discharge device  220 , such that cooled, expanded gaseous coolant exiting the turbine  218  is directed into the discharge device  220 . 
     The discharge device  220  may comprise at least one device or apparatus configured and positioned to receive, expand, and discharge (e.g., exhaust) the cooled, expanded gaseous coolant exiting the turbine  218 . As shown in  FIG. 2 , in some embodiments, the discharge device  220  comprises a secondary nozzle separate (e.g., discrete) from the combustor assembly  202 . In additional embodiments, the discharge device  220  comprises a component of the combustor assembly  202  configured and operated to receive and discharge the cooled, expanded gaseous fuel. For example, the discharge device  220  may comprise a portion of the nozzle  208 , such as a portion of the nozzle  208  configured and operated to facilitate supersonic flow therein. In such embodiments, the cooled, expanded gaseous fuel may be directed into the component (e.g., the nozzle  208 ) of the combustor assembly  202  without passing through and/or being acted upon (e.g., combusted) by one or more other components (e.g., the combustion chamber  206 ) of the combustor assembly  202 . The ability to exhaust the cooled, expanded gaseous coolant exiting the turbine  218  while combusting the pressurized liquid fuel directed into the combustor assembly  202  may permit the liquid-fueled rocket engine assembly  200  to achieve desirable operational parameters (e.g., relatively higher turbine power outputs, relatively elevated combustion chamber pressures, etc.) for a variety of different fuel types, coolant types, oxidizer types, fuel flowrates, coolant flowrates, and/or oxidizer flowrates without performance losses and/or failures that may otherwise be associated with configuring the liquid-fueled rocket engine assembly  200  to recirculate the cooled, expanded gaseous coolant. 
     During use and operation of the liquid-fueled rocket engine assembly  200 , a liquid fuel (e.g., liquid hydrogen; liquid ammonia; a liquid hydrocarbon, such as liquid methane, liquid propane, liquid kerosene, etc.; combinations thereof; etc.) may be directed from the fuel source  212  and into the fuel pump  216 . Pressurized liquid fuel may then exit the fuel pump  216  and may be directed into the injector  204  of the combustor assembly  202 . Within the injector  204 , the pressurized liquid fuel may be combined with pressurized liquid oxidizer (e.g., pressurized liquid oxygen, pressurized liquid hydrogen peroxide, combinations thereof, etc.) delivered from the oxidizer source  210  using the oxidizer pump  214  to form a reactant mixture. The reactant mixture may then be directed into and combusted within the combustion chamber  206  of the combustor assembly  202  to generate propellant gases, which may then be exhausted from the nozzle  208  of the combustor assembly  202  to generate thrust. Concurrently, a liquid coolant (e.g., liquid hydrogen, liquid ammonia, liquid methane, liquid water, etc.) may be directed from the coolant source  213  and into the coolant pump  217 . Pressurized liquid coolant may then exit the coolant pump  217  and may be directed into the combustor jacket  226 . Within the combustor jacket  226 , the pressurized liquid coolant may extract heat from the combustor assembly  202  to form a heated, gaseous coolant. The heated, gaseous coolant may then be directed into the turbine  218  to drive the turbine  218  and power the oxidizer pump  214 , the fuel pump  216 , and the coolant pump  217 . Cooled, expanded gaseous coolant may then exit the turbine  218  and may be directed into the discharge device  220 , where it may be expanded and exhausted. 
     Employing a liquid coolant (e.g., the liquid coolant from the coolant source  213 ) fluidly isolated from a liquid fuel (e.g., the liquid fuel from the fuel source  212 ) to facilitate and maintain an expander cycle (e.g., an open expander cycle) of the liquid-fueled rocket engine assembly  200  may reduce the complexity of the liquid-fueled rocket engine assembly  200  and may increase the number of liquid fuels that may be employed by the liquid-fueled rocket engine assembly  200  (e.g., by the combustor assembly  202  of the liquid-fueled rocket engine assembly  200  to generate thrust) as compared to conventional expander-cycle liquid-fueled rocket engine assemblies. For example, employing a separate liquid coolant to maintain an open expander cycle of the liquid-fueled rocket engine assembly  200  may eliminate the need for additional equipment (e.g., heat exchangers, mixers, etc.) that may otherwise be needed to maintain a closed expander cycle, and/or may permit the use of liquid fuels (e.g., higher molecular weight liquid fuels and/or lower specific heat liquid fuels, such as liquid propane, liquid kerosene, etc.) that may be unsuitable to maintain an expander cycle. 
     The liquid-fueled rocket engine assemblies (e.g., the liquid-fueled rocket engine assemblies  100 ,  200 ) and methods of the disclosure may provide improved performance (e.g., improved specific impulse), increased efficiency, increased reliability, reduced costs (e.g., material costs, equipment costs, etc.), reduced weight, increased simplicity, and/or increased safety as compared to many conventional liquid-fueled rocket engine assemblies and associated conventional methods. 
     While the disclosure is susceptible to various modifications and alternative forms, specific embodiments have been shown by way of example in the drawings and have been described in detail herein. However, the disclosure is not limited to the particular forms disclosed. Rather, the disclosure is to cover all modifications, equivalents, and alternatives falling within the scope of the disclosure as defined by the following appended claims and their legal equivalents.