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Gradl, Paul R. - One of the best experts on this subject based on the ideXlab platform.

  • Extreme-Temperature Carbon- and Ceramic-Matrix Composite Nozzle Extensions for Liquid Rocket Engines
    2019
    Co-Authors: Gradl, Paul R., Valentine, Peter G.
    Abstract:

    The United States (US) National Aeronautics and Space Administration (NASA) and its US industry partners are developing extreme-temperature composite nozzle extensions for a variety of cryogenic Liquid Rocket Engine propulsion systems. Applications under consideration at the NASA Marshall Space Flight Center (MSFC) include launch vehicle (upper stage), in-space, and lunar lander descent/ascent propulsion systems. Composite material systems of interest include carbon-carbon (C-C) and carbon/silicon-carbide (C-SiC) composites, as well as modified versions of such composites (coatings, mixed matrices, oxidation inhibitors, etc.). Missions addressed include access to low Earth orbit, the lunar surface, and more distant destinations. Cryogenic upper-stage and in-space Liquid Rocket Engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures, increasing exhaust velocities. As a result of the large size of such nozzles, and the related Engine performance requirements, composite nozzle extensions are being considered to reduce mass impacts. Currently, metallic and foreign composite nozzle extensions used on the Atlas V, Delta IV, Falcon 9, and Ariane 5 launch vehicles represent the state-of-the-art. Such extensions are limited to use at or slightly above 2000F (1093C). Materials under development by MSFC and its industry partners have the potential to operate at temperatures up to (or above) 4250F (2343C). Marshall Space Flight Center efforts are aimed at (a) further developing the technology and databases needed to enable the use of composite nozzle extensions on cryogenic Liquid Rocket Engines, (b) developing and demonstrating low-cost capabilities for testing and qualifying such nozzle extensions, and (c) advancing the US domestic supply chain for nozzle extensions. Through the Small Business Innovation Research (SBIR) and Small Business Technology Transfer (STTR) Programs, low-level internal MSFC research projects, and teaming arrangements with domestic industry partners, composite material design, development, hot-fire test, and evaluation efforts are progressing. Hot- fire Engine tests of subscale hardware have been conducted using oxygen/hydrogen (LOX/H2), oxygen/methane (LOX/CH4), and oxygen/kerosene (LOX/RP-1) propellants. These Engine tests at MSFC have enabled the evaluation of both heritage and state-of-the-art material systems, demonstrating the initial capabilities of the extreme temperature materials and their fabrication methods. Recent, ongoing, and potential future work supporting cryogenic propulsion systems development will be presented. The composite nozzle extension technology and test capabilities being developed are intended to support both NASA and US Department of Defense requirements, as well as those of the broader Commercial Space industry

  • Progress in Additively Manufactured Copper-Alloy GRCop-84, GRCop-42, and Bimetallic Combustion Chambers for Liquid Rocket Engines
    2019
    Co-Authors: Greene, Sandy E., Gradl, Paul R., Protz, Christopher S., Ellis, David L.
    Abstract:

    Additive Manufacturing (AM) has significantly evolved over the last decade for use in the aerospace industry, particularly for Liquid Rocket Engines. AM offers a considerable departure from traditional manufacturing to rapidly fabricate components with complex internal features. High performance Liquid Rocket Engine combustion chambers that operate in a high heat flux environment are fabricated using a copper-alloy liner with a series of integral coolant channels. Copper-alloys provide the necessary conductivity and material strength for adequate design margins offering high performance without the need for film coolant. Copper-alloys present unique challenges to properly melt the powder in laser-based AM processes due to their high reflectivity and conductivity. Starting in 2014, NASAs Marshall Space Flight Center (MSFC) and Glenn Research Center (GRC) have developed a process for AM of GRCop (Copper- Chrome-Niobium) alloys using Selective Laser Melting (SLM). GRCop, originally developed at GRC, is a high conductivity, high-strength, dispersion strengthened copper-alloy for use in high-temperature, high heat flux applications. NASA has completed significant material characterization and testing, along with hot-fire testing, to demonstrate that GRCop-42 and GRCop-84 alloys are suitable for use in combustion chambers. Additional development and testing has been completed on AM bimetallic chambers using GRCop-84 liners and superalloy jackets, fabricated using two Directed Energy Deposition (DED) processes: Electron Beam Freeform Fabrication (EBF) and Blown Powder DED. NASA completed hot-fire testing on various AM chambers using GRCop-84, GRCop-42, and bimetallic chambers in Liquid Oxygen (LOX)/Hydrogen, LOX/Methane, and LOX/Kerosene propellants

  • GRCop-42 Development and Hot-fire Testing Using Additive Manufacturing Powder Bed Fusion for Channel-Cooled Combustion Chambers
    2019
    Co-Authors: Ellis David, Protz Chris, Evans Laura, Cooper Ken, Garcia Chance, Gradl, Paul R.
    Abstract:

    GRCop-42 is a high conductivity, high-strength dispersion strengthened copper-alloy for use in high heat flux applications such as Liquid Rocket Engine combustion devices. This alloy is part of the family of NASA-developed GRCop, copper-chrome-niobium alloys. GRCop alloys were developed for harsh environments specific to regeneratively-cooled combustion chambers and nozzles with good oxidation resistance. Significant development was completed on the GRCop-84 and GRCop-42 alloys in the extruded wrought form demonstrating feasibility for combustion chambers. NASA has recently developed a process for additive manufacturing, specifically Powder Bed Fusion (PBF) or Selective Laser Melting (SLM), of GRCop-42 to establish parameters, characterize the material, and complete testing of components with complex internal features. This evolution of the GRCop-42 was based on the successful predecessor development work using GRCop-84 with the motivation of establishing a new copper-alloy option for use in NASA, government, and industry programs with SLM. A few advantages have been shown with the GRCop-42 that include higher conductivity and faster build speeds over the GRCop-84, and a simplified powder supply chain. Initial property development has shown that it is possible to produce high density builds with strengths equivalent to wrought GRCop-42 and a conductivity greater than GRCop-84. The GRCop-42 has completed process development and initial properties have been established. Several demonstrator combustion chambers have also been fabricated with the SLM GRCop-42 that include integral channels and closeouts. Additional test units have been fabricated and are completing substantial hot-fire testing to demonstrate performance of the material, process, and design

  • Channel Wall Nozzle Manufacturing and Hot-Fire Testing using a Laser Wire Deposition Closeout Technique for Liquid Rocket Engines
    2018
    Co-Authors: Greene, Sandy E., Brandsmeier Will, Gradl, Paul R.
    Abstract:

    A regeneratively-cooled nozzle for Liquid Rocket Engine applications is a significant cost of the overall Engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. NASA has been investigating and advancing methods for fabrication of Liquid Rocket Engine channel wall nozzles to realize further cost and schedule improvements. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA Marshall Space Flight Center (MSFC) completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Deposition Closeout (LWDC). This process has been developed to significantly reduce time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication and hot-fire testing performance is discussed in this paper

  • Channel Wall Nozzle Manufacturing and Hot-Fire Testing Using a Laser Wire Direct Closeout Technique for Liquid Rocket Engines
    2018
    Co-Authors: Brandsmeier Will, Greene, Sandy E., Gradl, Paul R.
    Abstract:

    A regeneratively-cooled nozzle for Liquid Rocket Engine applications is a significant cost of the overall Engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. The National Aeronautics and Space Administration (NASA) has been investigating and advancing methods for fabrication of Liquid Rocket Engine channel wall nozzles to realize further cost and schedule improvements over traditional techniques. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA's Marshall Space Flight Center (MSFC) has completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Direct Closeout (LWDC). This process has been developed to significantly reduce the time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication, and hot-fire testing are discussed in this paper

Greene, Sandy E. - One of the best experts on this subject based on the ideXlab platform.

  • Progress in Additively Manufactured Copper-Alloy GRCop-84, GRCop-42, and Bimetallic Combustion Chambers for Liquid Rocket Engines
    2019
    Co-Authors: Greene, Sandy E., Gradl, Paul R., Protz, Christopher S., Ellis, David L.
    Abstract:

    Additive Manufacturing (AM) has significantly evolved over the last decade for use in the aerospace industry, particularly for Liquid Rocket Engines. AM offers a considerable departure from traditional manufacturing to rapidly fabricate components with complex internal features. High performance Liquid Rocket Engine combustion chambers that operate in a high heat flux environment are fabricated using a copper-alloy liner with a series of integral coolant channels. Copper-alloys provide the necessary conductivity and material strength for adequate design margins offering high performance without the need for film coolant. Copper-alloys present unique challenges to properly melt the powder in laser-based AM processes due to their high reflectivity and conductivity. Starting in 2014, NASAs Marshall Space Flight Center (MSFC) and Glenn Research Center (GRC) have developed a process for AM of GRCop (Copper- Chrome-Niobium) alloys using Selective Laser Melting (SLM). GRCop, originally developed at GRC, is a high conductivity, high-strength, dispersion strengthened copper-alloy for use in high-temperature, high heat flux applications. NASA has completed significant material characterization and testing, along with hot-fire testing, to demonstrate that GRCop-42 and GRCop-84 alloys are suitable for use in combustion chambers. Additional development and testing has been completed on AM bimetallic chambers using GRCop-84 liners and superalloy jackets, fabricated using two Directed Energy Deposition (DED) processes: Electron Beam Freeform Fabrication (EBF) and Blown Powder DED. NASA completed hot-fire testing on various AM chambers using GRCop-84, GRCop-42, and bimetallic chambers in Liquid Oxygen (LOX)/Hydrogen, LOX/Methane, and LOX/Kerosene propellants

  • Channel Wall Nozzle Manufacturing and Hot-Fire Testing using a Laser Wire Deposition Closeout Technique for Liquid Rocket Engines
    2018
    Co-Authors: Greene, Sandy E., Brandsmeier Will, Gradl, Paul R.
    Abstract:

    A regeneratively-cooled nozzle for Liquid Rocket Engine applications is a significant cost of the overall Engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. NASA has been investigating and advancing methods for fabrication of Liquid Rocket Engine channel wall nozzles to realize further cost and schedule improvements. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA Marshall Space Flight Center (MSFC) completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Deposition Closeout (LWDC). This process has been developed to significantly reduce time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication and hot-fire testing performance is discussed in this paper

  • Channel Wall Nozzle Manufacturing and Hot-Fire Testing Using a Laser Wire Direct Closeout Technique for Liquid Rocket Engines
    2018
    Co-Authors: Brandsmeier Will, Greene, Sandy E., Gradl, Paul R.
    Abstract:

    A regeneratively-cooled nozzle for Liquid Rocket Engine applications is a significant cost of the overall Engine due to the complexities of manufacturing a large thin-walled structure that must operate in extreme temperature and pressure environments. The National Aeronautics and Space Administration (NASA) has been investigating and advancing methods for fabrication of Liquid Rocket Engine channel wall nozzles to realize further cost and schedule improvements over traditional techniques. The methods being evaluated are targeting increased scale required for current NASA and commercial space programs. Several advanced rapid fabrication methods are being investigated for forming of the inner liner, producing the coolant channels, closeout of the coolant channels, and fabrication of the manifolds. NASA's Marshall Space Flight Center (MSFC) has completed process development and subscale hot-fire testing of a series of these advanced fabrication channel wall nozzle technologies to gather performance data in a relevant environment. The primary fabrication technique being discussed in this paper is Laser Wire Direct Closeout (LWDC). This process has been developed to significantly reduce the time required for closeouts of regeneratively-cooled slotted liners. It allows for channel closeout to be formed in place in addition to the structural jacket without the need for channel fillers or complex tooling. Additional technologies were also tested as part of this program including water jet milling and arc-based additive manufacturing deposition. Each nozzle included different fabrication features, materials, and methods to demonstrate durability in a hot-fire environment. The results of design, fabrication, and hot-fire testing are discussed in this paper

Marco Pizzarelli - One of the best experts on this subject based on the ideXlab platform.

  • pseudo boiling and heat transfer deterioration while heating supercritical Liquid Rocket Engine propellants
    Journal of Supercritical Fluids, 2021
    Co-Authors: Francesco Nasuti, Marco Pizzarelli
    Abstract:

    Abstract Heating of Liquid propellants used as the coolant in Rocket Engines may lead to undesired phenomena such as pseudo-boiling or heat transfer deterioration under specific conditions. This can be an issue for propellants characterized at the same time by relatively low critical pressure and temperature. Light hydrocarbons, as for instance methane, belong to this family. In the present paper, a critical review is made of the main results obtained by Authors and their coworkers for the present application. Focus is on the correlations and trends inferred by their numerical simulations mainly carried out considering methane as the coolant, perhaps the most challenging one.

Gradl Paul - One of the best experts on this subject based on the ideXlab platform.

  • Additive Manufacturing of NASA HR-1 Material for Liquid Rocket Engine Component Applications
    2019
    Co-Authors: Protz Chris, Ellis David, Evans Laura, Po Chen, Katsarelis Colton, Jones Zachary, Gradl Paul
    Abstract:

    NASA HR-1 is a high-strength Fe-Ni superalloy designed to resist high pressure, hydrogen environment embrittlement, oxidation, and corrosion. NASA HR-1 was originally developed at NASA in the 1990s and derived from JBK-75 to increase strength and ductility in high-pressure hydrogen environments. The NASA HR-1 chemistry was formulated to meet requirements for Liquid Rocket Engine applications, specifically components used in a high-pressure hydrogen environment. Recent developments using additive manufacturing (AM) have made this material an attractive option for channel-cooled nozzles under the Rapid Analysis and Manufacturing Propulsion Technology (RAMPT) program and other Liquid Rocket Engine component applications. The RAMPT program has baselined to fully evolve and characterize NASA HR-1 material. NASA HR-1 meets materials requirements for Liquid Rocket Engine components, including good hydrogen resistance, high conductivity, good low cycle fatigue performance, and high elongation and strength for channel-cooled nozzles in high heat flux environments. Initial development and characterization has been completed using blown powder Directed Energy Deposition (DED) and Laser Powder Bed Fusion (L-PBF) additive manufacturing techniques to develop material test samples and nozzle hardware. NASA HR-1 powder has been sourced and characterized from several powder suppliers, and a series of development and hardware samples completed fabrication using DED and L-PBF. Characterization of the material has included heat treatment development, metallography, chemistry evaluations, mechanical testing, measurement of thermophysical properties, and fabrication of relevant nozzle hardware to demonstrate feasibility. This paper presents results from the process and early materials development and provide future development work including hardware fabrication

  • Additive Manufacturing and Hot-fire Testing of Bimetallic GRCop-84 and C-18150 Channel-Cooled Combustion Chambers using Powder Bed Fusion and Inconel 625 Hybrid Directed Energy Deposition
    2019
    Co-Authors: Zagorski Kevin, Protz Chris, Doshi Vishal, Mccallum Hannah, Gradl Paul
    Abstract:

    Additive manufacturing (AM) is an advanced fabrication technique that is demonstrating tremendous potential to reduce fabrication lead times and costs for Liquid Rocket Engine components. The additive manufacturing technology lends itself to fabricate components with complex features such as internal coolant channels in combustion chambers that would otherwise require complex manufacturing operations. A requirement for high performance Engines is to use high conductivity, high strength materials such as copper-alloys for combustion chamber liners to provide adequate wall temperatures and meet subsequent structural margins. A further requirement of this configuration is to minimize weight by defining and fabricating material in discrete locations as required. NASA and Industry partner, Virgin Orbit, have been working to advance these technologies through development of bimetallic additive manufacturing techniques under a public-private partnership through NASAs Announcement of Collaborative Opportunity (ACO). This partnership is advancing a bimetallic hybrid additively manufactured combustion chamber that integrates Powder Bed Fusion (PBF), specifically Selective Laser Melting (SLM), and Directed Energy Deposition (DED) blown powder techniques to optimize the chamber materials and subsequent assembly. The SLM process is being developed for the combustion chamber liner to use copper-alloys GRCop-84 (Copper-Chrome-Niobium) or C-18150 (Copper-Chrome-Zirconium). The hybrid DED blown powder technology is used to apply an integrated structural jacket and manifolds using an Inconel 625 superalloy on the outer surface of the SLM copper liner. The hybrid DED technology being used on this program is a DMG Mori Seiki AM machining center which integrates the DED blown powder with an integral subtractive (traditional) machining to minimize overall setups. A series of chambers were fabricated using these techniques with GRCop-84/Inconel 625 and C-18150/Inconel and hot-fire tested at NASA Marshall Space Flight Center (MSFC) in LOX/Kerosene (RP-1). This paper describes the process development to integrate these AM technologies into an integrated bimetallic assembly, the design of the chamber, results from hot-fire testing, and further development

  • Hybrid Additive Manufacturing Deposition and Selective Laser Melting Techniques Applied to Copper-Alloy Liquid Rocket Engine Combustion Chambers
    2018
    Co-Authors: Duggleby Andrew, Gradl Paul, Zagorski Kevin, Doshi Vishal
    Abstract:

    Virgin Orbit and the NASA Marshall Space Flight Center have partnered to fabricate a small, multi-metallic, regeneratively cooled thrust chamber by leveraging the strengths of two different types of additive manufacturing: Direct Metal Laser Sintering (DMLS) for the copper alloy liner and blown powder Directed Energy Deposition (DED) for the Inconel structural jacket. The DED is being developed using Virgin Orbit's DMG Mori Seiki hybrid additive/subtractive machining center to further enable unique processing and further cost savings. The materials chosen are preferred for high performance thrust chamber applications, representing a significant advancement from the compromises typically made in the production of metal 3D printed thrust chambers for Rocket Engines. The 1.2K-lbf thrust article is a 16" tall, 3" diameter design modified for this effort to be regeneratively cooled with water and hot fired using a RP-1/LOX pintle-style development injector from Virgin Orbit's NewtonFour upper stage Engine. The thrust chamber was tested at Virgin Orbit's Necker test site in Mojave, California. Results from this test campaign are reported in addition to information characterizing the liner, jacket, and bimetallic diffusion layer materials. This paper will also highlight some of the future bimetallic thrust chamber developments that MSFC and Virgin Orbit will complete under a recently awarded NASA contract

Annafederica Urbano - One of the best experts on this subject based on the ideXlab platform.

  • exploration of combustion instability triggering using large eddy simulation of a multiple injector Liquid Rocket Engine
    Combustion and Flame, 2016
    Co-Authors: Laurent Selle, Gabriel Staffelbach, Annafederica Urbano, Sebastien Ducruix, Benedicte Cuenot, Thomas Schmitt, Sebastien Candel
    Abstract:

    This article explores the possibility of analyzing combustion instabilities in liq- uid Rocket Engines by making use of Large Eddy Simulations (LES). Calculations are carried out for a complete small-scale Rocket Engine, including the injection manifold thrust chamber and nozzle outlet. The Engine comprises 42 coaxial injectors feeding the combustion chamber with gaseous hydrogen and Liquid oxy- gen and it operates at supercritical pressures with a maximum thermal power of 80 MW. The objective of the study is to predict the occurrence of transverse high-frequency combustion instabilities by comparing two operating points fea- turing different levels of acoustic activity. The LES compares favorably with the experiment for the stable load point and exhibits a nonlinearly unstable trans- verse mode for the experimentally unstable operating condition. A detailed analysis of the instability retrieves the experimental data in terms of spectral features. It is also found that modifications of the flame structures and of the global combustion region configuration have similarities with those observed in recent model scale experiments. It is shown that the overall acoustic activity mainly results from the combination of one transverse and one radial mode of the chamber, which are also strongly coupled with the oxidizer injectors.