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NASA NTRS2000

Lead-Free Double-Base Propellant for the 2.75 Inch Rocket Motor

The current MK 66 2.75 inch Rocket Motor double-base propellant contains the lead-based ballistic modifier LC-12-15 to achieve the desired plateau and mesa burning rate characteristics. The use of lead compounds poses a concern for the environment and for personal safety due to the metal's toxic nature when introduced into the atmosphere by propellant manufacture, rocket motor firing, and disposal. Copper beta-resorcylate (copper 2,4-di-hydroxy-benzoate) was successfully used in propellant as a simple modifier in the mid 1970's. This and other compounds have also been mixed with lead salts to obtain more beneficial ballistic results. Synthesized complexes of lead and copper compounds soon replaced the mixtures. The complexes incorporate the lead, copper lack of organic liquids, which allows for easier propellant processing. About ten years ago, the Indian Head Division, Naval Surface Warfare Center (NSWC), initiated an effort to develop a lead-free propellant for use in missile systems. Several lead-free propellant candidate formulations were developed. About five years ago, NSWC, in conjunction with Alliant Techsystems, Radford Army Ammunition Plant, continued ballistic modifier investigations. A four component ballistic modifier system without lead for double-base propellants that provide adequate plateau and mesa burn rate characteristics was developed and patented. The ballistic modifier's system contains bismuth subsalicylate, 1.5 percent; copper salicylate, 1.0 percent, copper stannate, 0.77 percent; and carbon black, 0.1 percent. Action time and impulse data obtained through multiple static firings indicate that the new lead-free double-base propellant, while not a match for NOSIH-AA-2, will be a very suitable replacement in the 2.75 inch Rocket Motor. Accelerated aging of the double-base propellant containing the lead-free ballistic modifier showed that it had a much higher rate of stabilizer depletion than the AA-2. A comprehensive study showed that an increased rate of stabilizer depletion occurred in propellants containing monobasic copper salicylate. The study also showed that propellants containing a mixture of bismuth subsalicylate and copper salicylate, had only about one-half the stabilizer depletion rate than those with copper salicylate alone. The copper salicylate catalyzes the decomposition of nitroglycerin, which triggers a chain of events leading to the increased rate of stabilizer depletion. A program has been initiated to coat the ballistic modifier, thus isolating it from the nitroglycerin.

Magill, B. T.↗

NASA NTRS1994

Handbook for Dynamic Data Acquisition and Analysis - IES Recommended Practices 012.1

The original motivation for this Handbook came from personnel of The Aerospace Corporation (TAC), who are technical advisors to the Air Force Space and Missile Systems Center. Over the preceding two decades, TAC has observed a seemingly endless series of errors and anomalies in the acquisition and analysis of structural dynamic and aeroacoustic data.

Shock vibration acoustic noise↗

NASA NTRS2010

Liquid Rocket Engine Testing - Historical Lecture: Simulated Altitude Testing at AEDC

The span of history covered is from 1958 to the present. The outline of this lecture draws from historical examples of liquid propulsion testing done at AEDC primarily for NASA's Marshall Space Flight Center (NASA/MSFC) in the Saturn/Apollo Program and for USAF Space and Missile Systems dual-use customers. NASA has made dual use of Air Force launch vehicles, Test Ranges and Tracking Systems, and liquid rocket altitude test chambers / facilities. Examples are drawn from the Apollo/ Saturn vehicles and the testing of their liquid propulsion systems. Other examples are given to extend to the family of the current ELVs and Evolved ELVs (EELVs), in this case, primarily to their Upper Stages. The outline begins with tests of the XLR 99 Engine for the X-15 aircraft, tests for vehicle / engine induced environments during flight in the atmosphere and in Space, and vehicle staging at high altitude. The discussion is from the author's perspective and background in developmental testing.

Dougherty, N. S.↗

NASA NTRS2012

DebriSat - A Planned Laboratory-Based Satellite Impact Experiment for Breakup Fragment Characterization

DebriSat is a planned laboratory ]based satellite hypervelocity impact experiment. The goal of the project is to characterize the orbital debris that would be generated by a hypervelocity collision involving a modern satellite in low Earth orbit (LEO). The DebriSat project will update and expand upon the information obtained in the 1992 Satellite Orbital Debris Characterization Impact Test (SOCIT), which characterized the breakup of a 1960 's US Navy Transit satellite. There are three phases to this project: the design and fabrication of an engineering model representing a modern, 50-cm/50-kg class LEO satellite known as DebriSat; conduction of a laboratory-based hypervelocity impact to catastrophically break up the satellite; and characterization of the properties of breakup fragments down to 2 mm in size. The data obtained, including fragment size, area ]to ]mass ratio, density, shape, material composition, optical properties, and radar cross ]section distributions, will be used to supplement the DoD fs and NASA fs satellite breakup models to better describe the breakup outcome of a modern satellite. Updated breakup models will improve mission planning, environmental models, and event response. The DebriSat project is sponsored by the Air Force fs Space and Missile Systems Center and the NASA Orbital Debris Program Office. The design and fabrication of DebriSat is led by University of Florida with subject matter experts f support from The Aerospace Corporation. The major milestones of the project include the complete fabrication of DebriSat by September 2013, the hypervelocity impact of DebriSat at the Air Force fs Arnold Engineering Development Complex in early 2014, and fragment characterization and data analyses in late 2014.

Liou, J.-C.↗

NASA NTRS2014

New Laboratory-Based Satellite Impact Experiments for Breakup Fragment Characterization

A consortium consisting of the NASA Orbital Debris Program Office, U.S. Air Force's Space and Missile Systems Center, the Aerospace Corporation, and University of Florida is planning a series of hypervelocity impact experiments on mockup targets at the U.S. Air Force's Arnold Engineering Development Complex (AEDC) in early 2014. The target for the first experiment resembles a rocket upper stage whereas the target for the second experiment represents a typical 60‐cm/50‐kg class payload that incorporates modern spacecraft materials and components as well as exterior wrap of multi‐layer insulation and three solar panels. The projectile is designed with the maximum mass that AEDC's Range G two‐stage light gas gun can accelerate to an impact speed of 7 km/sec. The impact energy is expected to be close to 15 MJ to ensure catastrophic destruction of the target after the impact. Low density foam panels are installed inside the target chamber to slow down and soft‐catch the fragments for post‐impact processing. Diagnostic instruments, such as x‐ray and high speed optical cameras, will also be used to record the breakup process. The main goal of this "DebriSat" project is to characterize the physical properties, including size, mass, shape, and density distributions, of orbital debris that would be generated by a hypervelocity collision involving an upper stage or a modern satellite in the low Earth orbit environment. In addition, representative fragments will be selected for laboratory optical and radar measurements to allow for better interpretation of data obtained by telescope and radar observations. This paper will provide a preliminary report of the impact results and the plans to process, measure, and analyze the fragments.

Liou, J.-C.↗

NASA NTRS2015

Orbital Debris Assesment Tesing in the AEDC Range G

The space environment presents many hazards for satellites and spacecraft. One of the major hazards is hypervelocity impacts from uncontrolled man-made space debris. Arnold Engineering Development Complex (AEDC), The National Aeronautics and Space Administration (NASA), The United States Air Force Space and Missile Systems Center (SMC), the University of Florida, and The Aerospace Corporation configured a large ballistic range to perform a series of hypervelocity destructive impact tests in order to better understand the effects of space collisions. The test utilized AEDC's Range G light gas launcher, which is capable of firing projectiles up to 7 km/s. A non-functional full-scale representation of a modern satellite called the DebriSat was destroyed in the enclosed range enviroment. Several modifications to the range facility were made to ensure quality data was obtained from the impact events. The facility modifcations were intended to provide a high impact energy to target mass ratio (>200 J/g), a non-damaging method of debris collection, and an instrumentation suite capable of providing information on the physics of the entire imapct event.

Polk, Marshall↗

NASA NTRS2016

Science and Technology (S and T) Roadmap Collaboration between SMC, NASA, and Government Partners

National Security Space (NSS) presents multi-faceted S and T challenges. We must continually innovate enterprise and information management; provide decision support; develop advanced materials; enhance sensor technology; transform communication technology; develop advanced propulsion and resilient space architectures and capabilities; and enhance multiple additional S and T domains. These challenges are best met by leveraging advanced S and T research and technology development from a number of DoD agencies and civil agencies such as NASA. The authors of this paper have engaged in these activities since 2006 and over the past decade developed multiple strategic S and T relationships. This paper highlights the Office of the Space Missile Systems Center (SMC) Chief Scientist (SMC/ST) collaboration with the NASA Office of Chief Technologist (NASA OCT), which has multiple S and T activities that are relevant to NSS. In particular we discuss the development of the Technology Roadmaps that benefit both Civil Space and NSS. Our collaboration with NASA OCT has been of mutual benefit to multiple participants. Some of the other DoD components include the Defense Advanced Research Projects agency (DARPA), Air Force Research Laboratory (AFRL), Naval Research Laboratory (NRL), The USAF Office of Chief Scientist, the USAF Science Advisory Board (SAB), Space and Naval Warfare Systems Command (SPAWAR), and a number of other services and agencies. In addition, the human talent is a key enabler of advanced S and T activities; it is absolutely critical to have a strong supply of talent in the fields of Science Technology, Engineering, and Mathematics (STEM). Consequently, we continually collaborate with the USAF Institute of Technology (AFIT), other service academies and graduate schools, and other universities and colleges. This paper highlights the benefits that result from such strategic S and T partnerships and recommends a way forward that will continually build upon these achievements into the future.

Science Technology Collaboration↗

NASA NTRS2017

Characterization of Orbital Debris via Hyper-Velocity Laboratory-Based Tests

Existing DoD and NASA satellite breakup models are based on a key laboratory test, Satellite Orbital debris Characterization Impact Test (SOCIT), which has supported many applications and matched on-orbit events involving older satellite designs reasonably well over the years. In order to update and improve these models, the NASA Orbital Debris Program Office, in collaboration with the Air Force Space and Missile Systems Center, The Aerospace Corporation, and the University of Florida, replicated a hypervelocity impact using a mock-up satellite, DebriSat, in controlled laboratory conditions. DebriSat is representative of present-day LEO satellites, built with modern spacecraft materials and construction techniques. Fragments down to 2 mm in size will be characterized by their physical and derived properties. A subset of fragments will be further analyzed in laboratory radar and optical facilities to update the existing radar-based NASA Size Estimation Model (SEM) and develop a comparable optical-based SEM. A historical overview of the project, status of the characterization process, and plans for integrating the data into various models will be discussed herein.

Cowardin, Heather↗

NASA NTRS2016

Overview of Marshall Space Flight Center Activities for the Combustion Stability Tool Development Program

This presentation covers the overall scope, schedule, and activities associated with the NASA - Marshall Space Flight Center (MSFC) involvement with the Combustion Stability Tool Development (CSTD) program. The CSTD program is funded by the Air Force Space & Missile Systems Center; it is approximately two years in duration and; and it is sponsoring MSFC to: design, fabricate, & execute multi-element hardware testing, support Air Force Research Laboratory (AFRL) single element testing, and execute testing of a small-scale, multi-element combustion chamber. Specific MSFC Engineering Directorate involvement, per CSTD-sponsored task, will be outlined. This presentation serves a primer for the corresponding works that provide details of the technical work performed by individual groups within MSFC.

Kenny, R. J.↗

NASA NTRS2016

Hot-Fire Test Results of Liquid Oxygen/RP-2 Multi-Element Oxidizer-Rich Preburners

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. To supply the oxidizer-rich combustion products to the main injector of the integrated test article, existing subscale preburner injectors from a previous NASA-funded oxidizer-rich staged combustion engine development program were utilized. For the integrated test article, existing and newly designed and fabricated inter-connecting hot gas duct hardware were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. However, before one of the preburners was used in the integrated test article, it was first hot-fire tested at length to prove it could provide the hot exhaust gas mean temperature, thermal uniformity and combustion stability necessary to perform in the integrated test article experiment. This paper presents results from hot-fire testing of several preburner injectors in a representative combustion chamber with a sonic throat. Hydraulic, combustion performance, exhaust gas thermal uniformity, and combustion stability data are presented. Results from combustion stability modeling of these test results are described in a companion paper at this JANNAF conference, while hot-fire test results of the preburner injector in the integrated test article are described in another companion paper.

Protz, C. S.↗

NASA NTRS2016

Design and Fabrication of Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged Combustion Thrust Chamber Injectors

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. On the current project, several configurations of new main injectors were considered for the thrust chamber assembly of the integrated test article. All the injector elements were of the gas-centered swirl coaxial type, similar to those used on the Russian oxidizer-rich staged-combustion rocket engines. In such elements, oxidizer-rich combustion products from the preburner/turbine exhaust flow through a straight tube, and fuel exiting from the combustion chamber and nozzle regenerative cooling circuits is injected near the exit of the oxidizer tube through tangentially oriented orifices that impart a swirl motion such that the fuel flows along the wall of the oxidizer tube in a thin film. In some elements there is an orifice at the inlet to the oxidizer tube, and in some elements there is a sleeve or "shield" inside the oxidizer tube where the fuel enters. In the current project, several variations of element geometries were created, including element size (i.e., number of elements or pattern density), the distance from the exit of the sleeve to the injector face, the width of the gap between the oxidizer tube inner wall and the outer wall of the sleeve, and excluding the sleeve entirely. This paper discusses the design rationale for each of these element variations, including hydraulic, structural, thermal, combustion performance, and combustion stability considerations. This paper also discusses the fabrication and assembly of the injector components, including the injector body/interpropellant plate, the additive manufactured GRCop-84 faceplate, and the pieces that make up the injector elements including the oxidizer tube, an inlet to the oxidizer tube, and a facenut that includes the fuel tangential inlets and forms the initial recessed volume where oxidizer and fuel first interact. Hot-fire test results of these main injector designs in an integrated test article that includes an oxidizer-rich preburner are described in companion papers at this JANNAF meeting.

Garcia, C. P.↗

NASA NTRS2016

Combustion Dynamics and Stability Modeling of a Liquid Oxygen/RP-2 Oxygen-Rich Staged Combustion Preburner and Thrust Chamber Assembly with Gas-Centered Swirl Coaxial Injector Elements

The Combustion Stability Tool Development (CSTD) project, funded by the Air Force Space and Missile Systems Center, began in March 2015 supporting a renewed interest in the development of a liquid oxygen/hydrocarbon, oxygen-rich combustion engine. The project encompasses the design, assembly, and hot-fire testing of the NASA Marshall Space Flight Center 40-klbf Integrated Test Rig (MITR). The test rig models a staged-combustion configuration by combining an oxygen-rich preburner (ORPB), to generate hot gas, with a thrust chamber assembly (TCA) using gas-centered swirl coaxial injector elements. There are five separately designed interchangeable injectors in the TCA that each contain 19- or 27- injector elements. A companion paper in this JANNAF conference describes the design characteristics, rationale, and fabrication issues for all the injectors. The data acquired from a heavily instrumented rig encompasses several injectors, several operating points, and stability bomb tests. Another companion paper in this JANNAF conference describes this test program in detail. In this paper, dynamic data from the hot-fire testing is characterized and used to identify the responses in the ORPB and TCA. A brief review of damping metrics are discussed and applied as a measure of stability margin for damped acoustic modes. Chug and longitudinal combustion stability models and predictions are described which includes new dynamic models for compressible flow through an orifice and a modification to incorporate a third feed line for inclusion of the fuel-film coolant. Flow-acoustics finite element modeling is used to investigate the anticipated TCA acoustics, the effects of injector element length on stability margin, and the potential use of an ORPB orifice trip ring for improving longitudinal stability margin.

Casiano, M. J.↗

NASA NTRS2016

Hot-Fire Test Results of an Oxygen/RP-2 Multi-Element Oxidizer-Rich Staged-Combustion Integrated Test Article

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center was contracted to assemble and hot-fire test a multi-element integrated test article demonstrating combustion characteristics of an oxygen/hydrocarbon propellant oxidizer-rich staged-combustion engine thrust chamber. Such a test article simulates flow through the main injectors of oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines, or future U.S.-built engine systems such as the Aerojet-Rocketdyne AR-1 engine or the Hydrocarbon Boost program demonstration engine. For the thrust chamber assembly of the test article, several configurations of new main injectors, using relatively conventional gas-centered swirl coaxial injector elements, were designed and fabricated. The design and fabrication of these main injectors are described in a companion paper at this JANNAF meeting. New ablative combustion chambers were fabricated based on hardware previously used at NASA for testing at similar size and pressure. An existing oxygen/RP-1 oxidizer-rich subscale preburner injector from a previous NASA-funded program, along with existing and new inter-connecting hot gas duct hardware, were used to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Results from independent hot-fire tests of the preburner injector in a combustion chamber with a sonic throat are described in companion papers at this JANNAF conference. The resulting integrated test article - which includes the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was assembled at Test Stand 116 at the East Test Area of the NASA Marshall Space Flight Center. The test article was well instrumented with static and dynamic pressure, temperature, and acceleration sensors to allow the collected data to be used for combustion analysis model development. Hot-fire testing was conducted with main combustion chamber pressures ranging from 1400 to 2100 psia, and main combustion chamber mixture ratios ranging from 2.4 to 2.9. Different levels of fuel film cooling injected from the injector face were examined ranging from none to about 12% of the total fuel flow. This paper presents the hot-fire test results of the integrated test article. Combustion performance, stability, thermal, and compatibility characteristics of both the preburner and the thrust chamber are described. Another companion paper at this JANNAF meeting includes additional and more detailed test data regarding the combustion dynamics and stability characteristics.

Hulka, J. R.↗

NASA NTRS2017

Characterization of Orbital Debris via Hyper-Velocity Laboratory-Based Tests

Existing DOD and NASA satellite breakup models are based on a key laboratory test, Satellite Orbital debris Characterization Impact Test (SOCIT), which has supported many applications and matched on-orbit events involving older satellite designs reasonably well over the years. In order to update and improve these models, the NASA Orbital Debris Program Office, in collaboration with the Air Force Space and Missile Systems Center, The Aerospace Corporation, and the University of Florida, replicated a hypervelocity impact using a mock-up satellite, DebriSat, in controlled laboratory conditions. DebriSat is representative of present-day LEO satellites, built with modern spacecraft materials and construction techniques. Fragments down to 2 mm in size will be characterized by their physical and derived properties. A subset of fragments will be further analyzed in laboratory radar and optical facilities to update the existing radar-based NASA Size Estimation Model (SEM) and develop a comparable optical-based SEM. A historical overview of the project, status of the characterization process, and plans for integrating the data into various models will be discussed herein.

Cowardin, Heather↗

NASA NTRS2018

Performance, Stability and Compatibility of Oxygen/RP-1 Multi-Element Oxidizer-Rich Staged-Combustion Injectors

In 2015 and 2016, the National Aeronautics and Space Administration Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/RP-1 propellant multi-element oxidizer-rich staged-combustion test article. The main objective was to provide thrust chamber combustion stability data as part of the Combustion Stability Tool Development program, although demonstration of performance and compatibility of oxidizer-rich main injectors was also important. Funding was provided by the Air Force Space and Missile Systems Center. Five configurations of main injectors were designed and fabricated, using conventional gas-centered swirl coaxial injector element designs generally similar to those used in oxygen/kerosene oxidizer-rich staged combustion engines such as the Russian RD-180 or NK-33 engines. Variations of element features included element size, recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. Ablative combustion chambers were fabricated based on hardware previously used at the NASA MSFC for testing at similar size and pressure. Existing oxygen/RP-1 oxidizer-rich subscale preburner injectors and hot gas ducts from a previous NASA-funded program were modified for use to supply the oxidizer-rich combustion products to the oxidizer circuit of the main injector of the thrust chamber. Testing of the resulting integrated test article - which included the preburner, inter-connecting hot gas duct, main injector, and ablative combustion chamber - was conducted at Test Stand 116 at the East Test Area of the NASA MSFC. The test article was well instrumented with static and dynamic pressure, temperature, and vibration sensors. This paper presents and discusses all the hot-fire test results of the integrated test article thrust chamber. Eighteen successful hot-fire tests of the integrated rig were conducted. Testing was accomplished with all five of the injector element concepts. Main combustion chamber pressures ranged from 710 to 2350 psia, and main combustion chamber mixture ratios ranged from 2.47 to 2.87. A chamber barrier fuel film coolant of about 2% to 4% of the total fuel flow was used for most tests. Characteristic exhaust velocity efficiency excluding the influence of the fuel film cooling ranged from 91% to 98% of theoretical. All tests of the thrust chamber exhibited stable combustion, even down to 40% of nominal operating pressures. Compatibility of the injector face and combustion chamber walls was acceptable. This paper is a follow-on to publication of preliminary test data presented at the 2016 JANNAF Liquid Propulsion Subcommittee meeting.

Hulka, J. R.↗

NASA NTRS2018

Comparison of Single-Element and Multi-Element Oxygen/RP-1 Oxidizer-Rich Staged-Combustion Injector Hot-Fire Test Results

As part of the Combustion Stability Tool Development project funded by the Air Force Space and Missile Systems Center, the NASA Marshall Space Flight Center designed, fabricated, assembled and hot-fire tested an oxygen/hydrocarbon propellant multi-element integrated test article that included an oxidizer-rich oxygen/hydrocarbon propellant preburner and a staged-combustion main injector. Also as part of this project, the Air Force Research Laboratory fabricated single-element main injectors of the same designs as used in the NASA multi-element injectors, and tested them in a staged-combustion integrated test article that used an oxidizer-rich oxygen/hydrogen propellant preburner. Final results of the multi-element and single-element staged-combustion main injector test programs are described in companion papers at this JANNAF meeting. The design, development, and preliminary test results of these main injectors have also been described in previous JANNAF papers. The main injector element designs were all based on relatively conventional gas-centered swirl coaxial injector element configurations such as used in Russian RD-170 and NK-33 engines, and planned for use in future U.S.-built experimental engine systems such as the Hydrocarbon Boost program demonstration engine. Four different elements were tested in both the multi-element and single-element main injectors, at similar combustion chamber pressures, chamber contraction ratios, and mixture ratios. Variations of the element features included recess depth, fuel gap width, and the presence of the sleeve separating the swirling fuel flow from the axial oxidizer flow. This paper compares the hydraulics, combustion performance, stability, and compatibility characteristics of the single-element and multi-element injectors operated at similar conditions. The single-element hardware is shown to have captured a significant level of the operability of the multi-element hardware.

Hulka, J. R.↗

NASA NTRS2019

Analysis of the DebriSat Fragments and Comparison to the NASA Standard Satellite Breakup Model

Existing DOD and NASA satellite breakup models are based on a key laboratory test, the 1992 Satellite Orbital debris Characterization Impact Test, which has supported many applications and matched on-orbit events involving older satellite designs reasonably well over the years. In 2014, the NASA Orbital Debris Program Office, in collaboration with the Air Force Space and Missile Systems Center, The Aerospace Corporation, and the University of Florida, conducted a hypervelocity impact test using a high-fidelity, mock-up satellite, DebriSat, in controlled and instrumented laboratory conditions to update and improve these models. DebriSat is representative of present-day, low Earth orbit satellites, having been constructed with modern spacecraft materials and techniques. The DebriSat fragment ensemble provided a variety of shapes, bulk densities, and dimensions. Fragments down to 2 mm in size are being characterized by their physical and derived properties. The data will inform updates to the current NASA Standard Satellite Breakup Model (SSBM), which was formulated using laboratory and ground-based measurements of on-orbit fragmentation events to describe an average breakup for spacecraft and upper-stage collisions and explosions. Although individual fragment collection and characterization is still ongoing, this paper will present the status of the DebriSat fragment data analysis including cumulative characteristic length and cumulative mass distributions, area-to-mass distributions, and characteristic length versus mass distributions. Additionally, comparisons to the NASA SSBM will be presented.

Murray, James↗

NASA NTRSDate not supplied

Progress in Red vs. Green Propellant Operations: ASCENT Load Operations Lessons Learned

The Green Propellant Infusion Mission (GPIM) [1] is a NASA Space Technology Mission Directorate (STMD) effort to establish via demonstration ASCENT propellant (Advanced Space Craft Energetic Non-Toxic propellant, formerly referred to as AF-M315E) ground operations through on –orbit performance. ASCENT was developed by Air Force Research Laboratory (AFRL) as an increased energy density, reduced toxicity and reduced hazard monopropellant formulation relative to hydrazine. The GPIM team was led by Ball Aerospace and Technologies Corporation (BATC) as Program Management and Principal Investigator with AFRL, Aerojet-Rocketdyne (AR), the United States Space Force Space and Missile Systems Center (USSF/SMC), NASA Glenn Research Center (GRC), NASA Goddard Space Flight Center (GSFC), and NASA Kennedy Space Center (KSC) as Co-Investigators. The AFRL co-investigator role in the GPIM program was to develop the Propellant Ground Support Equipment (PGSE), associated Standard Operating Procedures (SOP) and Check Lists (CL), propellant load, propellant specification standard conventions and analysis methods, decontamination and waste disposal practices. This paper will discuss lessons learned in documentation, crew composition and readiness, equipment readiness, propellant readiness and potential improvements in the PGSE design.

Anthony P Zuttarelli↗