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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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The analysis performed on the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) inlet housings is summarized. Three DIAL finite element models were build to aid in assessing the structural life of the welds and fillets at the vanes. Complete results are given.
A previously developed technique allows an estimate of integral mixing to be obtained from an image of laser scattered light from particle seeded fuel in the hypervelocity flow through a scramjet combustor. This previous mixing analysis formulation contains an assumption of a constant velocity flowfield across the plane of the fuel plume image. For high-speed scramjet combustors, the velocity flowfield is quite uniform and an assumption of constant velocity works well. Applying this same mixing analysis technique to fuel plume images obtained from a mid-speed scramjet combustor makes it desirable to remove the constant velocity assumption. This is due to the non-uniform velocity flowfields present in mid-speed scramjet combustors. A new formulation of the mixing analysis methodology is developed and presented so that the technique can be applied to a mid-speed scramjet combustor without the need to assume a constant velocity flowfield.
Results of a thermal/structural design analysis study of a fuel injection strut for an airframe integrated hydrogen cooled scramjet are presented. It is indicated that a feasible thermal/structural concept has been identified for the static load conditions and that thermal stresses dominate the response. It is suggested that the response of the concept to dynamic loads be investigated.
It is expected that future NASA mission energy storage requirements can be adequately met by regenerative fuel cells (RFCs). Aerospace RFCs are likely to package proton exchange membrane electrochemical conversion devices based on hydrogen, oxygen, and water. To optimize system designs and direct further development, the most useful system metric is round-trip efficiency (RTE). Improving RTE generally increases specific energy and reduces system mass. The following analysis models the impact of several factors on RTE for RFCs scaling from 0.1 to 50 kW. The analysis demonstrates that higher electrolyzer (EZ) operational pressure most negatively impacts both RFC RTE and specific energy. Reactant storage heating and power management losses are significant for RFCs of any scale, but solenoid valve power becomes a noticeable parasitic load for smaller RFCs. Spherical gas storage vessels benefit RTE by only 2 to 4 percentage points but increase specific energy by ~50 percent. Cryogenic reactant storage is shown to depress RTE to below 15 percent in all cases. For a given RFC scale, there is an RTE benefit to specifying larger EZs and recharging at higher rates—if the RFC can be coupled with a corresponding power source. Although fuel cell (FC) operation is less impactful on RTE than EZ factors, efficiency is maximized by sizing the FCs relatively large and keeping current density low. RFC design is a balance of many interrelated factors, but this analysis provides a starting point for difficult decisions that must be made when designing a lunar equator RFC.
It is expected that future NASA mission energy storage needs can be adequately met by regenerative fuel cells (RFCs). Aerospace RFCs package proton exchange membrane (PEM) electrochemical conversion devices based on hydrogen, oxygen, and water. To optimize system designs and direct further development, the most useful system metric is round-trip efficiency (RTE). Improving RTE generally increases specific energy and reduces system mass. The following analysis models the impact of several factors on RTE for RFCs at the lunar south pole scaling from 0.1 kW to 50 kW. The analysis demonstrates that higher electrolyzer operational pressure most negatively impacts both RFC RTE and specific energy. Optimal electrolyzer membrane thickness is evaluated as a function of operational current density and pressure. Reactant storage heating and power management losses are significant for RFCs of any scale but are reduced in impact at the south pole compared to colder equator applications. Solenoid valve power remains a noticeable parasitic load for smaller RFCs. Larger electrolyzers enable higher peak theoretical efficiencies when recharging at higher rates, if the RFC can be coupled with a corresponding higher power source. RFC design is a compromise of many interrelated factors, but this analysis provides a starting point for difficult decisions that must be made when designing a lunar south pole RFC.
This volume summarizes the analysis used to assess the structural life of the Space Shuttle Main Engine (SSME) High Pressure Fuel Turbo-Pump (HPFTP) Third Stage Impeller. This analysis was performed in three phases, all using the DIAL finite element code. The first phase was a static stress analysis to determine the mean (non-varying) stress and static margin of safety for the part. The loads involved were steady state pressure and centrifugal force due to spinning. The second phase of the analysis was a modal survey to determine the vibrational modes and natural frequencies of the impeller. The third phase was a dynamic response analysis to determine the alternating component of the stress due to time varying pressure impulses at the outlet (diffuser) side of the impeller. The results of the three phases of the analysis show that the Third Stage Impeller operates very near the upper limits of its capability at full power level (FPL) loading. The static loading alone creates stresses in some areas of the shroud which exceed the yield point of the material. Additional cyclic loading due to the dynamic force could lead to a significant reduction in the life of this part. The cyclic stresses determined in the dynamic response phase of this study are based on an assumption regarding the magnitude of the forcing function.
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A computational fluid dynamics model with porosity and permeability formulations in the transport equations has been developed to study the concept of nuclear thermal propulsion through the analysis of a pulsed irradiation of a particle bed element (PIPE). The numerical model is a time-accurate pressure-based formulation. An adaptive upwind scheme is employed for spatial discretization. The upwind scheme is based on second- and fourth-order central differencing with adaptive artificial dissipation. Multiblocked porosity regions have been formulated to model the cold frit, particle bed, and hot frit. Multiblocked permeability regions have been formulated to describe the flow shaping effect from the thickness-varying cold frit. Computational results for several zero-power density PIPEs and an elevated-particle-temperature PIPE are presented. The implications of the computational results are discussed.
A propulsion system analysis of the dual fuel, dual throat engine for launch vehicle applications was conducted. Basic dual throat engine characterization data were obtained to allow vehicle optimization studies to be conducted. A preliminary baseline engine system was defined.
A materials compatibility and thermal stability investigation was conducted using five common liquid hydrocarbon fuels and two structural materials. The tests were performed at the NASA Glenn Research Center Heated Tube Facility under environmental conditions similar to those encountered in regeneratively cooled rocket engines. Scanning-electron microscopic analysis in conjunction with energy dispersive spectroscopy (EDS) was utilized to characterize the condition of the tube inner wall surface and any carbon deposition or corrosion that was formed during selected runs. Results show that the carbon deposition process in stainless steel tubes was relatively insensitive to fuel type or test condition. The deposition rates were comparable for all fuels and none of the stainless steel test pieces showed any signs of corrosion. For tests conducted with copper tubing, the sulfur content of the fuel had a significant impact on both the condition of the tube wall and carbon deposition rates. Carbon deposition rates for the lowest sulfur fuels (2 ppm) were slightly higher than those recorded in the stainless steel tubes with no corrosion observed on the inner wall surface. For slightly higher sulfur content (25 ppm) fuels, nodules that intruded into the flow area were observed to form on the inner wall surface. These nodules induced moderate tube pressure drop increases. The highest sulfur content fuels (400 ppm) produced extensive wall pitting and dendritic copper sulfide growth that was continuous along the entire tube wall surface. The result of this tube degradation was the inability to maintain flow rate due to rapidly increasing test section pressure drops. Accompanying this corrosion were carbon deposition rates an order of magnitude greater than those observed in comparable stainless steel tests. The results of this investigation indicate that trace impurities in fuels (i.e. sulfur) can significantly impact the carbon deposition process and produce unacceptable corrosion levels in copper based structural materials.
Failure analysis of tooling boss weld in lower bulkhead of Saturn 1C stage fuel tank
A computational fluid dynamics model which simulates the steady state operation of the SSME fuel preburner is developed. Specifically, the model will be used to quantify the flow factors which cause local hot spots in the fuel preburner in order to recommend experiments whereby the control of undesirable flow features can be demonstrated. The results of a two year effort to model the preburner are presented. In this effort, investigating the fuel preburner flowfield, the appropriate transport equations were numerically solved for both an axisymmetric and a three-dimensional configuration. Continuum's VAST (Variational Solution of the Transport equations) code, in conjunction with the CM-1000 Engineering Analysis Workstation and the NASA/Ames CYBER 205, was used to perform the required calculations. It is concluded that the preburner operational anomalies are not due to steady state phenomena and must, therefore, be related to transient operational procedures.
CFD analysis is presented of the mixing characteristics and performance of three fuel injectors at hypervelocity flow conditions. The calculations were carried out using the VULCAN-CFD solver and Reynolds-Averaged Simulations (RAS). The high Mach number flow conditions match those proposed for the planned experiments conducted as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. The EIMP aims to investigate scramjet fuel injection and mixing physics, improve the understanding of underlying physical processes, and develop enhancement strategies and functional relationships relevant to flight Mach numbers greater than eight. Because of the high Mach number flow considered, the injectors consist of a fuel placement device, a strut; and a fluidic vortical mixer, a ramp. These devices accomplish the necessary task of distributing and mixing fuel into the supersonic cross-flow albeit via different strategies. Both of these devices were previously studied at lower flight Mach numbers where they exhibited promising performance in terms of mixing efficiency and total pressure recovery. For comparison, a flush-wall injector is also included. This type of injector generally represents the simplest method of introducing fuel into a scramjet combustor, however, at high flight Mach number conditions, the dynamic pressure needed to induce sufficient fuel penetration may be difficult to achieve along with other requirements such as achieving desired levels of fuel-to-air mixing at the required equivalence ratio. The three injectors represent the baseline configurations planned for the experiments. The current work discusses the mixing flow field behavior and differences among the three fuel injectors, mixing performance as described by the mixing efficiency and the total pressure recovery, and performance considerations based on the thrust potential.
Feasibility, parametric design, and comparative analysis of unguided solid fuel rocket vehicle with solar orbit, earth probe, and reentry mission capabilities
The problems of combustion instability in an annular combustion chamber are investigated. A modified Galerkin method was used to produce a set of modal amplitude equations from the general nonlinear partial differential acoustic wave equation. From these modal amplitude equations, the two variable perturbation method was used to develop a set of approximate equations of a given order of magnitude. These equations were modeled to show the effects of velocity sensitive combustion instabilities by evaluating the effects of certain parameters in the given set of equations. By evaluating these effects, parameters which cause instabilities to occur in the combustion chamber can be ascertained. It is assumed that in the annular combustion chamber, the liquid propellants are injected uniformly across the injector face, the combustion processes are distributed throughout the combustion chamber, and that no time delay occurs in the combustion processes.
High frequency combustion instability problems in a liquid fuel annular combustion chamber are examined. A modified Galerkin method was used to produce a set of modal amplitude equations from the general nonlinear partial differential acoustic wave equation in order to analyze the problem of instability. From these modal amplitude equations, the two variable perturbation method was used to develop a set of approximate equations of a given order of magnitude. These equations were modeled to show the effects of velocity sensitive combustion instabilities by evaluating the effects of certain parameters in the given set of equations.
The speed, range, payload, and fuel efficiency of a general aviation airplane powered by one turboprop engine was determined and compared to a twin engine turboprop aircraft. An airplane configuration was developed which can carry six people for a noreserve range of 2,408 km at a cruise speed above 154 m/s, and a cruise altitude of about 9,144 m. The cruise speed is comparable to that of the fastest of the current twin turboprop powered airplanes. It is found that the airplane has a cruise specific range greater than all twin turboprop engine airplanes flying in its speed range and most twin piston engine airplanes flying at considerably slower cruise airspeeds.
Pulsed ruby laser sheet illumination of the spray is used for the initial recording of data on very-high-resolution photographic film. The digitization of mosaic elements is effected with a vidicon and video digitizer whose output is stored in computer RAM memory for processing. Highly nonspherical elements and a broad range of drop diameters (8-2000 microns) resulting from the unusual rheological properties of the fuel-additive system are accommodated by the device configuration and algorithms. It is found that the generation of two-dimensional images by means of scattered light also eliminates errors resulting from variations in the index of refraction and from the submicron scattering sites that are often present within the modified fuel. No a priori information on the drop size distribution or on the system response to various drop sizes is required.