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All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Centrifugal Nuclear Thermal Rocket Challenges and Potential

The Centrifugal Nuclear Thermal Rocket (CNTR) is a liquid fueled fission propulsion concept designed to heat propellant to 5000 K prior to expansion through a nozzle. A specific impulse up to 1800 s may be achieved using hydrogen propellant, and a specific impulse up to 1000 s may be achieved using more storable propellants such as methane, ammonia, or propane. The high uranium density of the liquid metallic uranium or liquid uranium carbide fuel will help enable compact engines suitable for missions such as fast (<15 month) round trip human Mars missions or high delta-V missions in cislunar space. Long term applications of the CNTR could include the advanced exploration and utilization of the solar system through direct use of in-situ volatiles as propellant. Challenges associated with the CNTR are numerous. Centrifugal force is used to retain the liquid fuel in rotating fuel cylinders, and rotational velocities up to 5000 rpm may be required. Propellant flow must be directed such that all structures and moderators in the core are adequately cooled prior to the propellant entering the liquid fuel and being heated to 5000 K. The rotating fuel cylinder wall (RFCW) must have an inner surface designed to be compatible with liquid uranium metal or uranium carbide fuel up to at least 1500 K, and that inner surface may also need to be textured to help maintain acceptable wall temperatures. Propellant must flow radially inward through the RFCW while fuel is simultaneously contained. The RFCW should ideal-ly be made from a material with low neutron absorption to help minimize engine mass and facilitate the use of High Assay Low Enriched Uranium (HALEU) fuel in the system. The drive system for the rotating fuel cylinders must support all phases of operation. This paper will discuss computational and experimental research being conducted to address some of the challenges associated with the CNTR, and will also note potential mission benefits from the CNTR.

Michael G. Houts↗
NASA NTRS2010

Documentation of Stainless Steel Lithium Circuit Test Section Design

The Early Flight Fission-Test Facilities (EFF-TF) team was tasked by Naval Reactors Prime Contract Team (NRPCT) to design, fabricate, and test an actively pumped lithium (Li) flow circuit. This Li circuit takes advantage of work in progress at the EFF TF on a stainless steel sodium/potassium (NaK) circuit. The effort involved modifying the original stainless steel NaK circuit such that it could be operated with Li in place of NaK. This new design considered freeze/thaw issues and required the addition of an expansion tank and expansion/extrusion volumes in the circuit plumbing. Instrumentation has been specified for Li and circuit heaters have been placed throughout the design to ensure adequate operational temperatures and no uncontrolled freezing of the Li. All major components have been designed and fabricated prior to circuit redesign for Li and were not modified. Basic circuit components include: reactor segment, Li to gas heat exchanger, electromagnetic liquid metal pump, load/drain reservoir, expansion reservoir, instrumentation, and trace heaters. The reactor segment, based on a Los Alamos National Laboratory 100-kW design study with 120 fuel pins, is the only prototypic component in the circuit. However, due to earlier funding constraints, a 37-pin partial-array of the core, including the central three rings of fuel pins (pin and flow path dimensions are the same as those in the full design), was selected for fabrication and test. This Technical Publication summarizes the design and integration of the pumped liquid metal Li flow circuit as of May 1, 2005.

Godfroy, T. J.↗
NASA NTRS2010

Documentation of Stainless Steel Lithium Circuit Test Section Design

The Early Flight Fission-Test Facilities (EFF-TF) team was tasked by Naval Reactors Prime Contract Team (NRPCT) to design, fabricate, and test an actively pumped lithium (Li) flow circuit. This Li circuit takes advantage of work in progress at the EFF TF on a stainless steel sodium/potassium (NaK) circuit. The effort involved modifying the original stainless steel NaK circuit such that it could be operated with Li in place of NaK. This new design considered freeze/thaw issues and required the addition of an expansion tank and expansion/extrusion volumes in the circuit plumbing. Instrumentation has been specified for Li and circuit heaters have been placed throughout the design to ensure adequate operational temperatures and no uncontrolled freezing of the Li. All major components have been designed and fabricated prior to circuit redesign for Li and were not modified. Basic circuit components include: reactor segment, Li to gas heat exchanger, electromagnetic liquid metal pump, load/drain reservoir, expansion reservoir, instrumentation, and trace heaters. The reactor segment, based on a Los Alamos National Laboratory 100-kW design study with 120 fuel pins, is the only prototypic component in the circuit. However, due to earlier funding constraints, a 37-pin partial-array of the core, including the central three rings of fuel pins (pin and flow path dimensions are the same as those in the full design), was selected for fabrication and test. This Technical Publication summarizes the design and integration of the pumped liquid metal Li flow circuit as of May 1, 2005. This supplement contains drawings, analysis, and calculations

Godfroy, Thomas J.↗
NASA NTRS1990

A computational design tool for scramjet combustor film cooling and fuel mixing predictions

A methodology for the analysis of a film cooled combustor of a hydrogen-fueled, supersonic combustion ramjet (scramjet) is presented. The SHIP3D parabolic flow code, based on an extension of the pressure-correction method of Patankar and Spalding for compressible flow is used to compute the flow field in the combustor. The effect of film cooling on the overall vehicle performance is then calculated using the one-dimensional, multistep cycle code, CPIPE. Results are presented for a conceptual combustor at Mach 16 flight conditions. It is seen that film cooling can have a significant effect on the engine performance.

Kamath, Pradeep S.↗
NASA NTRS1980

Some advantages of methane in an aircraft gas turbine

Because liquid methane may be obtained from existing natural gas sources or produced synthetically from a range of other hydrocarbon sources (coal, biomass, shale, organic waste), it is considered as an aviation fuel in a simplified cycle analysis of the performance of a turboprop engine intended for operation at Mach 0.8 and 10,688 m altitude. Performance comparisons are given for four cases in which the turbine cooling air is either not cooled or cooled to -111, -222, and -333 K, and the advantages and problems that may be expected from direct use of the cryogenic fuel in turbine cooling are discussed. It is shown that while (1) methane combustion characteristics are appreciably different from those of Jet A fuel and will require the development of different combustor designs, and (2) the safe integration of methane cryotanks into transport aircraft structures poses a major design problem, a highly fuel-efficient turboprop engine fueled by methane appears to be feasible.

Graham, R. W.↗
NASA NTRS1982

Energy Efficient Engine (E3) controls and accessories detail design report

An Energy Efficient Engine program has been established by NASA to develop technology for improving the energy efficiency of future commercial transport aircraft engines. As part of this program, a new turbofan engine was designed. This report describes the fuel and control system for this engine. The system design is based on many of the proven concepts and component designs used on the General Electric CF6 family of engines. One significant difference is the incorporation of digital electronic computation in place of the hydromechanical computation currently used.

Beitler, R. S.↗
NASA NTRS1993

Impeller flow field characterization with a laser two-focus velocimeter

Use of Computational Fluid Dynamics (CFD) codes, prevalent in the rocket engine turbomachinery industry, necessitates data of sufficient quality and quantity to benchmark computational codes. Existing data bases for typical rocket engine configurations, in particular impellers, are limited. In addition, traditional data acquisition methods have several limitations: typically transducer uncertainties are 0.5% of transducer full scale and traditional pressure probes are unable to provide flow characteristics in the circumferential (blade-to-blade) direction. Laser velocimetry circumvents these limitations by providing +0.5% uncertainty in flow velocity and +0.5% uncertainty in flow angle. The percent of uncertainty in flow velocity is based on the measured value, not full range capability. The laser electronics multiple partitioning capability allows data acquired between blades as the impeller rotates, to be analyzed separately, thus providing blade-to-blade flow characterization. Unlike some probes, the non-intrusive measurements made with the laser velocimeter does not disturb the flow. To this end,, and under Contract (NAS8-38864) to the National Aeronautics and Space Administration (NASA) at Marshall Space Flight Center (MSFC), an extensive test program was undertaken at Rocketdyne. Impellers from two different generic rocket engine pump configurations were examined. The impellers represent different spectrums of pump design: the Space Shuttle Main Engine (SSME) high pressure fuel turbopump (HPFTP) impeller was designed in the 1 1970's the Consortium for CFD application in Propulsion Technology Pump Stage Technology Team (Pump Consortium) optimized impeller was designed with the aid of modern computing techniques. The tester configuration for each of the impellers consisted of an axial inlet, an inducer, a diffuser, and a crossover discharge. While the tested configurations were carefully chosen to be representative of generic rocket engine pumps, several features of both testers were intentionally atypical. A crossover discharge, downstream of the impeller, rather than a volute discharge was used to minimize asymmetric flow conditions that might be reflected in the impeller discharge flow data. Impeller shroud wear ring radial clearances were purposely close to minimize leakage flow, thus increasing confidence in using the inlet data as an input to CFD programs. The empirical study extensively examined the flow fields of the two impellers via performance of laser two-focus velocimeter surveys in an axial plane upstream of the impellers and in multiple radial planes downstream of the impellers. Both studies were performed at the impeller design flow coefficients. Inlet laser surveys that provide CFD code inlet boundary conditions were performed in one axial plane, with ten radial locations surveyed. Three wall static pressures, positioned circumferentially around the impeller inlet, were used to identify asymmetrical pressure distributions in the inlet survey plane. The impeller discharge flow characterization consisted of three radial planes for the SSME HPFTP impeller and two radial planes for the Pump Consortium optimized impeller. Housing wall static pressures were placed to correspond to the radial locations surveyed with the laser velocimeter. Between five and thirteen axial stations across the discharge channel width were examined in each radial plane during the extensive flow mapping. The largely successful empirical flow characterization of two different impellers resulted in a substantial contribution to the limited existing data base, and yielded accurate data for CFD code benchmarking.

Brozowski, L. A.↗
NASA NTRS2016

Conceptual Design of a Single-Aisle Turboelectric Commercial Transport With Fuselage Boundary Layer Ingestion

A single-aisle commercial transport concept with a turboelectric propulsion system architecture was developed assuming entry into service in 2035 and compared to a similar technology conventional configuration. The turboelectric architecture consisted of two underwing turbofans with generators extracting power from the fan shaft and sending it to a rear fuselage, axisymmetric, boundary layer ingesting fan. Results indicate that the turbo- electric concept has an economic mission fuel burn reduction of 7%, and a design mission fuel burn reduction of 12% compared to the conventional configuration. An exploration of the design space was performed to better understand how the turboelectric architecture changes the design space, and system sensitivities were run to determine the sensitivity of thrust specific fuel consumption at top of climb and propulsion system weight to the motor power, fan pressure ratio, and electrical transmission efficiency of the aft boundary layer ingesting fan.

Welstead, Jason R.↗
NASA NTRS1984

Energy Efficient Engine combustor test hardware detailed design report

The Energy Efficient Engine (E3) Combustor Development effort was conducted as part of the overall NASA/GE E3 Program. This effort included the selection of an advanced double-annular combustion system design. The primary intent was to evolve a design which meets the stringent emissions and life goals of the E3 as well as all of the usual performance requirements of combustion systems for modern turbofan engines. Numerous detailed design studies were conducted to define the features of the combustion system design. Development test hardware was fabricated, and an extensive testing effort was undertaken to evaluate the combustion system subcomponents in order to verify and refine the design. Technology derived from this development effort will be incorporated into the engine combustion system hardware design. This advanced engine combustion system will then be evaluated in component testing to verify the design intent. What is evolving from this development effort is an advanced combustion system capable of satisfying all of the combustion system design objectives and requirements of the E3. Fuel nozzle, diffuser, starting, and emissions design studies are discussed.

Burrus, D. L.↗
NASA NTRS2019

Pilot Injector Redesign to Reduce N+3 Cycle Emissions for a Gas-Turbine Combustor

An overview is given of an effort for the use of CFD analysis to complement design and configuration definition of third generation Lean-Direct Injection combustion concepts (LDI-3) for NASA’s N+3 program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for a three-cup, nineteen-element flame tube array with redesigned pilot injectors to improve spray and emissions characteristics when compared to a previous LDI-3 design. All computations were performed with a consistent approach to mesh-generation, spray modeling, ignition and kinetics modeling for a ‘medium-power’ cycle condition. Computational predictions of the aerodynamics of a new pre-filming pilot injector were used to arrive at an optimized aerothermal design that meets effective area and fuel-air mixing criteria. The newly designed pilot injectors were shown to provide considerable improvements in aerodynamic stability, flame-tube pattern factor and NOx emissions, when compared to the original design.

nitrogen oxides injectors↗
NASA NTRS2019

Pilot Injector Redesign to Reduce N+3 Cycle Emissions for a Gas-Turbine Combustor

An overview is given of an effort for the use of CFD analysis to complement design and configuration definition of third generation Lean-Direct Injection combustion concepts (LDI-3) for NASA's N plus 3 program. The National Combustion Code (OpenNCC) was used to perform non-reacting and two-phase reacting flow computations for a three-cup, nineteen-element flametube array with redesigned pilot injectors to improve spray and emissions characteristics when compared to a previous LDI-3 design. All computations were performed with a consistent approach to mesh-generation, spray modeling, ignition and kinetics modeling for a "medium-power" cycle condition. Computational predictions of the aerodynamics of a new pre-filming pilot injector were used to arrive at an optimized aerothermal design that meets effective area and fuel-air mixing criteria. The newly designed pilot injectors were shown to provide considerable improvements in aerodynamic stability, flame-tube pattern factor and NOx emissions, when compared to the original design.

Nitrogen Oxides↗
NASA NTRS2006

Designing capture trajectories to unstable periodic orbits around Europa

The hostile environment of third body perturbations restricts a mission designer's ability to find well-behaved reproducible capture trajectories when dealing with limited control authority as is typical with low-thrust missions. The approach outlined in this paper confronts this shortcoming by utilizing dynamical systems theory and an extensive preexisting database of Restricted Three Body Problem (RTBP) periodic orbits. The stable manifolds of unstable periodic orbits are utilized to attract a spacecraft towards Europa. By selecting an appropriate periodic orbit, a mission designer can control important characteristics of the captured state including stability, minimum altitudes, characteristic inclinations, and characteristic radii among others. Several free parameters are optimized in the non-trivial mapping from the RTBP to a more realistic model. Although the ephemeris capture orbit is ballistic by design, low-thrust is used to target the state that leads to the capture orbit, control the spacecraft after arriving on the unstable quasi-periodic orbit, and begin the spiral down towards the science orbit. The approach allows a mission designer to directly target fuel efficient captures at Europa in an ephemeris model. Furthermore, it provides structure and controllability to the design of capture trajectories that reside in a chaotic environment.

low-thrust↗
NASA NTRS1981

Cell module and fuel conditioner development

A phosphoric acid fuel cell (PAFC) stack design having a 10 kW power rating for operation at higher than atmospheric pressure based on the existing Mark II design configuration is described. Functional analysis, trade studies and thermodynamic cycle analysis for requirements definition and system operating parameter selection purposes were performed. Fuel cell materials and components, and performance testing and evaluation of the repeating electrode components were characterized. The state of the art manufacturing technology for all fuel cell components and the fabrication of short stacks of various sites were established. A 10 kW PAFC stack design for higher pressure operation utilizing the top down systems engineering aproach was developed.

Feret, J. M.↗
NASA NTRS2014

Conceptual Design Optimization of an Augmented Stability Aircraft Incorporating Dynamic Response Performance Constraints

This research focused on incorporating stability and control into a multidisciplinary de- sign optimization on a Boeing 737-class advanced concept called the D8.2b. A new method of evaluating the aircraft handling performance using quantitative evaluation of the sys- tem to disturbances, including perturbations, continuous turbulence, and discrete gusts, is presented. A multidisciplinary design optimization was performed using the D8.2b transport air- craft concept. The con guration was optimized for minimum fuel burn using a design range of 3,000 nautical miles. Optimization cases were run using xed tail volume coecients, static trim constraints, and static trim and dynamic response constraints. A Cessna 182T model was used to test the various dynamic analysis components, ensuring the analysis was behaving as expected. Results of the optimizations show that including stability and con- trol in the design process drastically alters the optimal design, indicating that stability and control should be included in conceptual design to avoid system level penalties later in the design process.

Welstead, Jason↗
NASA NTRSDate not supplied

Parallel Hybrid Turboprop Performance Modeling and Optimization

NASA’s Electrified Powertrain Flight Demonstration (EPFD) project conducts ground and flight tests of integrated Megawatt (MW) class hybrid-electric powertrain systems on regional turboprop aircraft demonstrators. To meet the increased demand for assessment of potential capabilities and benefits from these novel vehicle configurations, NASA is developing tooling and models to estimate the performance of hybridized regional turboprops. This paper covers the development of a parametrically driven performance model for a De Havilland Canada Dash 8-400 (Q400) regional turboprop integrated with a novel parallel hybrid architecture using the Gascon framework. Gascon is a modern reimplementation of the General Aviation Synthesis Program (GASP) built using the Condor mathematical modeling framework in Python. Within Gascon, a parametric representation of the parallel hybrid architecture was synthesized, which features the electric motor coupled to the power turbine. This capability allows for in-the-loop optimization of the parametric parallel hybrid architecture to characterize the mission capabilities and fuel savings of the design and determine optimal power scheduling strategies for efficient electric power management for a given mission. The study shows that a fuel savings of up to 20% can be achieved, but that increased fuel savings comes at the expense of payload capacity.

Gascon↗