A Scramjet/Ramjet Heat Exchanger Analysis Tool (SRHEAT™) has been developed for rapid analyses of complex thermal cooling systems. Thermal management is critical to the development of dual-mode scramjets for hypersonic aerospace propulsion, which have high thermal loading with limited availability of heat sink sources. It is necessary that rapid trade studies of the thermal management system be accomplished to optimize the system for weight and cooling efficiency. To meet this need, SPIRITECH has developed a scramjet/ramjet heat exchanger design and optimization tool that performs a thermal analysis of the heat exchanger, assesses its structural strength, and optimizes the heat exchanger design to minimize the cooling flow requirement and the heat exchanger weight. FuelDev™ has been developed as an add-on to SRHEAT™ to provide a system level tool allowing fuels developers to experiment with fuel databases and to perform “what if” scenarios to determine the system level impacts of changes in fuel properties. FuelDev™ provides a tool to compare and contrast existing fuels for use in hypersonic vehicle thermal management systems. This tool provides the user with the ability to quantify the sensitivity of the thermal management system to changes in fuel properties afforded by new fuel technologies. The detailed heat exchanger design features included in the tools (i.e. geometry, material properties, fuel/coolant properties, etc.) make SRHEAT™ and FuelDev™ a valuable suite of tools in scramjet and hypersonic vehicle development, providing the low cost analytical capabilities that make possible the efficient development of aerospace components and fuels.
A Material Development tool (MatDev™) has been developed as an add-on module to complement SPIRITECH Advanced Products, Inc’s existing Scramjet/Ramjet Heat Exchanger Analysis Tool (SRHEAT™). SRHEAT™ is a scramjet/ramjet heat exchanger design and optimization tool that performs a thermal analysis of a heat exchanger, assesses its structural integrity, and optimizes the heat exchanger design to minimize the cooling flow requirement and the heat exchanger weight. SRHEAT™ can be used to evaluate and design complex thermal cooling systems, like those found in dual-mode scramjets for hypersonic aerospace propulsion, that have high thermal loading with limited availability of heat sink sources. MatDev™, in combination with SRHEAT™, is a trade study tool that is used to compare and contrast high temperature metal alloys and high temperature composites (i.e. CMC, C/SiC) in fuel cooled, heat exchanger liner panel design applications. MatDev™ uses typical engineering stress calculations and formulations to assess non-traditional stresses, such as interlaminar tension, interlaminar shear, in-plane shear, and flexure. These non-traditional calculations allow MatDev™ to accurately evaluate laminated, directional materials, such as high temperature composites. MatDev™ provides detailed outputs summarizing the resulting stress, weight, optimum heat exchanger design, and cooling flow requirements to provide the user with critical insight into the key drivers of the heat exchanger system of interest.
A dynamic simulator is being developed to demonstrate all modes of supersonic operation, including mode transition, for a Turbine-Based Combined Cycle (TBCC) propulsion system. The High Mach Transient Engine Cycle Code (HiTECC) is a highly integrated simulation tool comprised of models for each of the TBCC systems whose performance and controllability affect the thrust and operability of the propulsion system. The reported work details the development of the Thermal Management and Fuel System model conducted in the second year of a multiyear effort to develop a dynamic TBCC simulator. Once completed, this model will significantly extend the state-of-the-art for all TBCC modes of operation by providing a numerical simulation of the systems, interactions, and transient responses affecting the ability of the propulsion system to transition from turbine-based to ramjet/scramjet-based propulsion.
A dynamic simulator is being developed to demonstrate all modes of operation, including mode transition, for a Turbine-Based Combined Cycle (TBCC) propulsion system. The High Mach Transient Engine Cycle Code (HiTECC) is a highly integrated simulation tool comprised of models for each of the TBCC systems whose performance and controllability affect the thrust and operability of the propulsion system. The reported work details the development of the Hydraulic and Kinematic System models conducted in the second year of a multiyear effort to develop a dynamic TBCC simulator. Once completed, this model will significantly extend the state-of-the-art for all TBCC modes of operation by providing a numerical simulation of the systems, interactions, and transient responses affecting the ability of the propulsion system to transition from turbine-based to ramjet/scramjet-based propulsion.
A Turbine-Based Combined Cycle (TBCC) dynamic simulation model is being developed to demonstrate all modes of operation, including mode transition, for a turbine-based combined cycle propulsion system. The High Mach Transient Engine Cycle Code (HiTECC) is a highly integrated tool comprised of modules for modeling each of the TBCC systems whose interactions and controllability affect the TBCC propulsion system thrust and operability during its modes of operation. By structuring the simulation modeling tools around the major TBCC functional modes of operation (Dry Turbojet, Afterburning Turbojet, Transition, and Dual Mode Scramjet) the TBCC mode transition and all necessary intermediate events over its entire mission may be developed, modeled, and validated. The reported work details the development of the gas turbine and dual-mode scramjet performance models conducted in the first year of a multiyear effort to develop a dynamic TBCC simulation model. Once completed, this model will significantly extend the state-of-the-art for all TBCC modes of operation by providing a numerical simulation of the systems, interactions, and transient responses affecting the ability of the propulsion system to transition from turbine-based to ramjet/scramjet-based propulsion while maintaining constant thrust.
The Scramjet/Ramjet Heat Exchanger Analysis Tool (SRHEAT) developed by SPIRITECH allows rapid analyses of the complex integrated cooling and structural systems required for hypersonic air-breathing propulsion systems. The ability to handle detailed heat exchanger and structural design features allows for near real time evaluation of performance and weight sensitivities to geometry, coolant path properties, flight conditions, construction materials, and fuels. These features make SRHEAT invaluable for initial assessment of hypersonic propulsion systems, providing the low cost analytical capability that makes rapid evaluation of design trade space possible. The ability to perform these trade studies as part of the early conceptual design phase is crucial to making correct initial design decisions before conducting more expensive preliminary and detail design efforts. These broad capabilities of SRHEAT are illustrated with a series of example trade studies, showing how this tool can be used to insure selection of the best combination of geometry, thermal and structural designs, sizing, vehicle integration, and mission flight path.
A Scramjet/Ramjet Heat Exchanger Analysis Tool (SRHEAT™) has been developed for rapid analyses of complex thermal cooling systems. The detailed heat exchanger design features included in this code (i.e. geometry, material properties, fuel/coolant properties, etc.) make SRHEAT™ a valuable tool in scramjet and hypersonic vehicle development, providing the low cost analytical capabilities that make possible the efficient development of aerospace components. A key feature of SRHEAT™ is its ability to optimize the heat exchanger thermal design for minimum fuel flow requirement while providing a structurally viable design. Optimization includes both the ordering of the cooling flow circuit and the sizing of the heat exchanger channels. Large computational times are required for standard optimization techniques due to the sheer number of interdependent variables associated with the complex thermal management system. Several methods have been developed and adapted to reduce computational time requirements of optimization. The result is a fast code with the built-in intelligence to make design decisions leading to an optimized thermal management system design.
A Scramjet/Ramjet Heat Exchanger Analysis Tool (SRHEAT™) has been developed for rapid analyses of complex thermal cooling systems. Thermal management is critical to the development of dual-mode scramjets for hypersonic aerospace propulsion, which have high thermal loading with limited availability of heat sink sources. It is necessary that rapid trade studies of the thermal management system be accomplished to optimize the system for weight and cooling efficiency. To meet this need, SPIRITECH has developed a scramjet/ramjet heat exchanger design and optimization tool that performs a thermal analysis of the heat exchanger, assesses its structural strength, and optimizes the heat exchanger design to minimize the cooling flow requirement and the heat exchanger weight. Radiation, conduction, and convection are all included to accurately model this complex aero/thermal system. The user can select the coolant/fuel from various jet fuels (with endothermic properties) or common combustible fluids (H2 & CH4). In addition, the option for several high temperature materials are included. The code is packaged with a user-friendly interface to simplify its use within large trade studies. The detailed heat exchanger design features included in the code (i.e. geometry, material properties, fuel/coolant properties, etc.) make SRHEAT™ a valuable tool in scramjet and hypersonic vehicle development, providing the low cost analytical capabilities that make possible the efficient development of aerospace components.
A user-friendly heat transfer/thermal modeling code, LinerTherm™, has been developed for rapid analysis of aircraft exhaust liner cooling systems. Most of the thermal modeling codes on the market today fall into one of two categories either they are complex 3-D codes requiring significant engineering resources or they are simple conduction models that lack advanced convection capabilities required for high performance aerospace and propulsion cooling applications. In the high-tech field of aerospace propulsion, it is necessary that rapid modeling of complex thermal systems be accomplished to enable the trade studies required to optimize aircraft cooling liner designs for weight, cost, and performance. This code performs detailed thermal analysis of gas turbine exhaust liners used in the augmentor and nozzle and includes cooling capabilities for impingement, multi-hole film-cooled, slot filmcooled, and convectively cooled liners. Its simple user interface provides the capability of performing quick trade studies for a vast array of cooling liner designs, including a wide selection of included materials.
A study has been completed to evaluate the merits of using injection of high-pressure air to control a hypersonic vehicle s pitching moment without adversely impacting the installed nozzle performance. A 3D CFD model was developed and used to investigate the feasibility of using fluidic injection for vehicle control. The underlying critical parameters necessary to control the shock wave location were defined and their effects were quantified. Results have shown that variations in the injection pressure and flow provide changes in the oblique shock angle and that the pressures acting on the SERN ramp are increased in the region of the shock impingement on the ramp. The increase in pressure results in a corresponding change in vehicle moment. However, results have also shown that the resulting performance, when calculated as CFGsec=F/(Fidp+Fids), decreased as flow was injected, providing a net system loss. The parameters that may be used to control the angle of the oblique shock wave are the pressure, flow, and angle of injection flow. The pressure and flow were independently controlled in the matrix of CFD runs analyzed. The effect of pressure and flow on oblique shock angle are combined in the ideal thrust of the injectant (secondary) flow, which was found to be a critical correlating parameter. Although it is understood that the injection angle is also a critical parameter, its effects were not investigated in detail in this study.
Advances in aircraft performance depend heavily upon improved and properly integrated propulsion systems. Historically, new engines and aircraft are developed concurrently, but the design and test cycle of engine systems is longer than that of the aircraft they power because of demanding flight qualification, reliability, and durability requirements1. Consequently, the engine hardware development process starts first, so that the success of the entire program often hinges on engine design decisions made early in the process. Critical to the design of efficient air vehicle systems is the design of the gas turbine exhaust nozzle. Aircraft exhaust nozzles serve two primary functions. First, they must control the engine backpressure to provide the correct, and optimum, engine performance, which is accomplished through jet area variations. Second, they must efficiently convert the potential energy of the exhausting gas to kinetic energy by increasing the exhaust velocity, which is accomplished through efficiently expanding the exhausting gases to the ambient pressure. Since the exhaust nozzle provides the integration between the propulsion and aircraft systems, its design must also consider installed, or thrust minus-drag, performance. Additional design challenges are introduced by the requirement for features such as thrust vectoring and reversing. This paper provides guidelines and procedures for incorporating these considerations into the design of gas turbine exhaust nozzles.
High transonic drag is an issue that must be considered in the design of supersonic flight vehicles. Generally, aftbody drag is maximum at flight Mach numbers near 1.0. This study has shown numerically that fluidic injection can be used to decrease nozzle aftbody drag, thereby increasing nozzle performance under transonic conditions. The fluidic injection is used to separate the flow expanding over the external flap, increasing the static pressure and decreasing the aftbody drag. The amount of flow injected was identified as a critical performance parameter while injection pressure was found to have only secondary effects. Use of multiple injection locations provided greater drag reduction for a given amount of injection flow. However, the location of these injectors must be considered.
A test was conducted to measure the time-averaged flow rate of a pulse detonation engine. The objective of this flow test was to determine the flow effectiveness of a pulse detonation engine utilizing a rotating spool of tubes. Since thrust is directly proportional to flow, the ability of the device to pass flow at operating rotational speeds is critical to its ability to create thrust. The flow effectiveness at the interface between the statically mounted tubes and the rotating spool of tubes is driven by the time-varying open area of the tubes and the flow coefficient at the interface. This test measured the flow coefficient for variations in the time-varying open area at the interface resulting from various tube diameters and rotational speeds. While the rotational speed of the “detonation” tubes did affect the flow effectiveness at the tube/tube interface, the effect was minimal and does not limit the feasibility of a pulse detonation engine.
