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Hulka, James

Publications and source records attributed to Hulka, James.

13 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2019

Uncertainty Analysis of Experimental Discharge Coefficients in Additively Manufactured Liquid Injector Elements

Screening of two additively manufactured liquid injector designs was conducted in the UAH high pressure spray facility. Four variants of each geometry with slightly different dimensions were obtained from eleven separate commercial additive manufacturing services. The devices were manufactured from Inconel 625 using the selective laser melting (SLM) powder bed process. The devices were cold flowed with water over a range of relevant pressure drops (75 psi to 1500 psi) to produce water flow rates from 0.037 to 1.75 lbm/s into ambient back pressure. Discharge coefficients determined from the testing along with the associated uncertainties provide insight into characteristic flow performance variabilities that can be expected from the SLM process for similar geometries.

Venters, James P.↗
NASA NTRS2018

Additive Manufacturing of Liquid Rocket Engine Combustion Devices: A Summary of Process Developments and Hot-Fire Testing Results

Additive Manufacturing (AM) is an emerging technology with a focus on complex metallic component fabrication: Enables complex shapes and internal features that were not possibly with traditional manufacturing techniques, and significant schedule and overall lifecycle cost reductions. To date at the NASA Marshall Space Flight Center (MSFC), combustion devices component hardware ranging in size from 100 - 35,000 lbf has been designed and manufactured using AM and many of these pieces have been hot-fire tested.

Gradl, Paul↗
NASA NTRS2018

Additive Manufacturing of Liquid Rocket Engine Combustion Devices: A Summary of Process Developments and Hot-Fire Testing Results

Additive Manufacturing (AM) of metals is a processing technology that has significantly matured over the last decade. For liquid propellant rocket engines, the advantages of AM for replacing conventional manufacturing of complicated and expensive metallic components and assemblies are very attractive. AM can significantly reduce hardware cost, shorten fabrication schedules, increase reliability by reducing the number of joints, and improve hardware performance by allowing fabrication of designs not feasible by conventional means. The NASA Marshall Space Flight Center (MSFC) has been involved with various forms of metallic additive manufacturing for use in liquid rocket engine component design, development, and testing since 2010. The AM technique most often used at the NASA MSFC has been powder-bed fusion or selective laser melting (SLM), although other techniques including laser directed energy deposition (DED), arc-based deposition, and laser-wire cladding techniques have also been used to develop several components. The purpose of this paper is to discuss the various internal programs at the NASA MSFC using AM to develop combustion devices hardware. To date at the NASA MSFC, combustion devices component hardware ranging in size from 100 lbf to 35,000 lbf have been designed and manufactured using SLM and deposition-based AM processes, and many of these pieces have been hot-fire tested. Combustion devices component hardware have included thrust chamber injectors, injector components such as faceplates, regeneratively-cooled combustion chambers, regeneratively-cooled nozzles, gas generator and preburner hardware, and augmented spark igniters. Ongoing and future developments for combustion devices have also included design of components sized for boost-class engines. Several design and hot-fire test iterations have been completed on these subscale and larger scale components, and a summary of these results will be presented as well.

Gradl, Paul↗
NASA NTRS2010

Throttling Characteristics of the RL10 Derivative Common Extensible Cryogenic Engine -- Demo 1.6 and 1.7 Test Results

The Common Extensible Cryogenic Engine (CECE) is a deep throttling cryogenic Lunar Module Descent Engine (LMDE) technology development demonstrator based on the Pratt & Whitney Rocketdyne (PWR) RL10 engine. Previous testing on this engine occurred during 2006 as Demo 1.0, then Demo 1.5 in 2007, Demo 1.6 in 2008 and finally Demo 1.7 in 2010. A review of Demo 1.0, 1.5, and preliminary results of 1.6 were reported in previous JANNAF papers. Demo 1.6 was tested at the PWR E-6 test facility in November 2008. The primary goal of this series was to mitigate low frequency combustion instability observed at low power levels. To mitigate the chug, the Demo 1.6 injector was modified from the previous configurations to include an approximately 0.050 inch thick teflon-type spray-on insulation to reduce heat transfer to the LOX manifold, which was believed to be a significant contributor to the low power instability. In addition, gaseous helium injection into the LOX manifold was used as a means to stabilize the system. Also explored in this test series was mitigation for a low power 1 Hz fuel system oscillation caused by sub-critical hydrogen boiling in the chamber cooling jacket. Reduced area gas venturis were utilized to avoid the 1 Hz fuel-size oscillation by keeping the cooling jacket supercritical down to lower engine power levels. Demo 1.7 began testing in March 2010. Its primary objectives were to demonstrate closed loop control capability on mixture ratio and chamber pressure, start to lower power levels and increased throttling ramp rates. Secondary test objectives that are discussed include multiple engine starts, higher mixture ratio excursions and additional time at low power level. While the complete test series and data reduction is not yet complete for Demo 1.7, an overview and up-to-date status is provided.

Devine, Matthew K.↗
NASA NTRS2009

Liquid-Propellant Rocket Engine Throttling: A Comprehensive Review

Liquid-Propellant Rocket Engines (LREs) are capable of on-command variable thrust or thrust modulation, an operability advantage that has been studied intermittently since the late 1930s. Throttleable LREs can be used for planetary entry and descent, space rendezvous, orbital maneuvering including orientation and stabilization in space, and hovering and hazard avoidance during planetary landing. Other applications have included control of aircraft rocket engines, limiting of vehicle acceleration or velocity using retrograde rockets, and ballistic missile defense trajectory control. Throttleable LREs can also continuously follow the most economical thrust curve in a given situation, compared to discrete throttling changes over a few select operating points. The effects of variable thrust on the mechanics and dynamics of an LRE as well as difficulties and issues surrounding the throttling process are important aspects of throttling behavior. This review provides a detailed survey of LRE throttling centered around engines from the United States. Several LRE throttling methods are discussed, including high-pressure-drop systems, dual-injector manifolds, gas injection, multiple chambers, pulse modulation, throat throttling, movable injector components, and hydrodynamically dissipative injectors. Concerns and issues surrounding each method are examined, and the advantages and shortcomings compared.

Casiano, Matthew↗
NASA NTRS2008

Scaling of Performance in Liquid Propellant Rocket Engine Combustors

The objectives are: a) Re-introduce to you the concept of scaling; b) Describe the scaling research conducted in the 1950s and early 1960s, and present some of their conclusions; c) Narrow the focus to scaling for performance of combustion devices for liquid propellant rocket engines; and d) Present some results of subscale to full-scale performance from historical programs. Scaling is "The ability to develop new combustion devices with predictable performance on the basis of test experience with old devices." Scaling can be used to develop combustion devices of any thrust size from any thrust size. Scaling is applied mostly to increase thrust. Objective is to use scaling as a development tool. - Move injector design from an "art" to a "science"

Hulka, James↗
NASA NTRS2006

Cold Flow Testing for Liquid Propellant Rocket Injector Scaling and Throttling

Scaling and throttling of combustion devices are important capabilities to demonstrate in development of liquid rocket engines for NASA's Space Exploration Mission. Scaling provides the ability to design new injectors and injection elements with predictable performance on the basis of test experience with existing injectors and elements, and could be a key aspect of future development programs. Throttling is the reduction of thrust with fixed designs and is a critical requirement in lunar and other planetary landing missions. A task in the Constellation University Institutes Program (CUIP) has been designed to evaluate spray characteristics when liquid propellant rocket engine injectors are scaled and throttled. The specific objectives of the present study are to characterize injection and primary atomization using cold flow simulations of the reacting sprays. These simulations can provide relevant information because the injection and primary atomization are believed to be the spray processes least affected by the propellant reaction. Cold flow studies also provide acceptable test conditions for a university environment. Three geometric scales - 1/4- scale, 1/2-scale, and full-scale - of two different injector element types - swirl coaxial and shear coaxial - will be designed, fabricated, and tested. A literature review is currently being conducted to revisit and compile the previous scaling documentation. Because it is simple to perform, throttling will also be examined in the present work by measuring primary atomization characteristics as the mass flow rate and pressure drop of the six injector element concepts are reduced, with corresponding changes in chamber backpressure. Simulants will include water and gaseous nitrogen, and an optically accessible chamber will be used for visual and laser-based diagnostics. The chamber will include curtain flow capability to repress recirculation, and additional gas injection to provide independent control of the backpressure. This paper provides a short review of the appropriate literature, as well as descriptions of plans for experimental hardware, test chamber instrumentation, diagnostics, and testing.

Kenny, Jeremy R.↗
NASA NTRS2006

Local Heat Flux Measurements with Single Element Coaxial Injectors

To support the mission for the NASA Vision for Space Exploration, the NASA Marshall Space Flight Center conducted a program in 2005 to improve the capability to predict local thermal compatibility and heat transfer in liquid propellant rocket engine combustion devices. The ultimate objective was to predict and hence reduce the local peak heat flux due to injector design, resulting in a significant improvement in overall engine reliability and durability. Such analyses are applicable to combustion devices in booster, upper stage, and in-space engines, as well as for small thrusters with few elements in the injector. In this program, single element and three-element injectors were hot-fire tested with liquid oxygen and ambient temperature gaseous hydrogen propellants at The Pennsylvania State University Cryogenic Combustor Laboratory from May to August 2005. Local heat fluxes were measured in a 1-inch internal diameter heat sink combustion chamber using Medtherm coaxial thermocouples and Gardon heat flux gauges. Injectors were tested with shear coaxial and swirl coaxial elements, including recessed, flush and scarfed oxidizer post configurations, and concentric and non-concentric fuel annuli. This paper includes general descriptions of the experimental hardware, instrumentation, and results of the hot-fire testing for three of the single element injectors - recessed-post shear coaxial with concentric fuel, flush-post swirl coaxial with concentric fuel, and scarfed-post swirl coaxial with concentric fuel. Detailed geometry and test results will be published elsewhere to provide well-defined data sets for injector development and model validatation.

Jones, Gregg↗
NASA NTRS2006

Local Heat Flux Measurements with Single and Small Multi-element Coaxial Element-Injectors

To support NASA's Vision for Space Exploration mission, the NASA Marshall Space Flight Center conducted a program in 2005 to improve the capability to predict local thermal compatibility and heat transfer in liquid propellant rocket engine combustion devices. The ultimate objective was to predict and hence reduce the local peak heat flux due to injector design, resulting in a significant improvement in overall engine reliability and durability. Such analyses are applicable to combustion devices in booster, upper stage, and in-space engines with regeneratively cooled chamber walls, as well as in small thrust chambers with few elements in the injector. In this program, single and three-element injectors were hot-fire tested with liquid oxygen and gaseous hydrogen propellants at The Pennsylvania State University Cryogenic Combustor Laboratory from May to August 2005. Local heat fluxes were measured in a 1-inch internal diameter heat sink combustion chamber using Medtherm coaxial thermocouples and Gardon heat flux gauges, Injector configurations were tested with both shear coaxial elements and swirl coaxial elements. Both a straight and a scarfed single element swirl injector were tested. This paper includes general descriptions of the experimental hardware, instrumentation, and results of the hot-fire testing for three coaxial shear and swirl elements. Detailed geometry and test results the for shear coax elements has already been published. Detailed test result for the remaining 6 swirl coax element for the will be published in a future JANNAF presentation to provide well-defined data sets for development and model validation.

Jones, Gregg↗
NASA NTRS2005

Status of the Combustion Devices Injector Technology Program at the NASA MSFC

To support the NASA Space Exploration Mission, an in-house program called Combustion Devices Injector Technology (CDIT) is being conducted at the NASA Marshall Space Flight Center (MSFC) for the fiscal year 2005. CDIT is focused on developing combustor technology and analysis tools to improve reliability and durability of upper-stage and in-space liquid propellant rocket engines. The three areas of focus include injector/chamber thermal compatibility, ignition, and combustion stability. In the compatibility and ignition areas, small-scale single- and multi-element hardware experiments will be conducted to demonstrate advanced technological concepts as well as to provide experimental data for validation of computational analysis tools. In addition, advanced analysis tools will be developed to eventually include 3-dimensional and multi- element effects and improve capability and validity to analyze heat transfer and ignition in large, multi-element injectors.

Jones, Gregg↗