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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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NASA NTRS2009

Apollo Spacecraft and Saturn V Launch Vehicle Pyrotechnics/Explosive Devices

The Apollo Mission employs more than 210 pyrotechnic devices per mission.These devices are either automatic of commanded from the Apollo spacecraft systems. All devices require high reliability and safety and most are classified as either crew safety critical or mission critical. Pyrotechnic devices have a wide variety of applications including: launch escape tower separation, separation rocket ignition, parachute deployment and release and electrical circuit opening and closing. This viewgraph presentation identifies critical performance, design requirements and safety measures used to ensure quality, reliability and performance of Apollo pyrotechnic/explosive devices. The major components and functions of a typical Apollo pyrotechnic/explosive device are listed and described (initiators, cartridge assemblies, detonators, core charges). The presentation also identifies the major locations and uses for the devices on: the Command and Service Module, Lunar Module and all stages of the launch vehicle.

Interbartolo, Michael↗
NASA NTRS2008

Estimating Shock Spectra: Extensions beyond GEVS

Shock response spectra (SRS) are the standard description of some vibration environments on spacecraft for equipment qualification. For shock events produced by pyrotechnic devices, SRS can have significant frequency content as high as 10 kHz. It is difficult to construct and analyze finite element models that can resolve dynamic behavior at such high frequencies. GEVS provides simple, empirically based methods for approximating the SRS for a wide variety of shock events. It begins with a base SRS according to the type of pyrotechnic device, and then provides attenuation relations to adjust this SRS according to distance from the shock source, the type of structural frame and the properties of any structural joints between the source and equipment. In our paper we extend GEVS to include more detailed information about the spacecraft structure. To retain the general framework of GEVS, we begin with a base SRS and adjust this SRS using attenuation relations. We use modal and traveling wave concepts to derive the attenuation relations for simple canonical structures. Then we show how these concepts can be used to analyze more complex structures using finite element mode shapes to explicitly calculate the attenuation factors. Since the low- to mid-frequency finite element modal information is extrapolated to obtain the low- to high-frequency attenuation relations, the resulting attenuated SRS is formulated as an upper bound rather than as mean predicted values. We illustrate the extended GEVS approach by analyzing the impact response of composite tubes and the shock response of the STEREO spacecraft.

Igusa, Takeru↗
NASA NTRS2009

Liquid Oxygen/Liquid Methane Test Results of the RS-18 Lunar Ascent Engine at Simulated Altitude Conditions at NASA White Sands Test Facility

Tests were conducted with the RS-18 rocket engine using liquid oxygen (LO2) and liquid methane (LCH4) propellants under simulated altitude conditions at NASA Johnson Space Center White Sands Test Facility (WSTF). This project is part of NASA's Propulsion and Cryogenics Advanced Development (PCAD) project. "Green" propellants, such as LO2/LCH4, offer savings in both performance and safety over equivalently sized hypergolic propulsion systems in spacecraft applications such as ascent engines or service module engines. Altitude simulation was achieved using the WSTF Large Altitude Simulation System, which provided altitude conditions equivalent up to ~122,000 ft (~37 km). For specific impulse calculations, engine thrust and propellant mass flow rates were measured. LO2 flow ranged from 5.9 - 9.5 lbm/sec (2.7 - 4.3 kg/sec), and LCH4 flow varied from 3.0 - 4.4 lbm/sec (1.4 - 2.0 kg/sec) during the RS-18 hot-fire test series. Propellant flow rate was measured using a coriolis mass-flow meter and compared with a serial turbine-style flow meter. Results showed a significant performance measurement difference during ignition startup due to two-phase flow effects. Subsequent cold-flow testing demonstrated that the propellant manifolds must be adequately flushed in order for the coriolis flow meters to give accurate data. The coriolis flow meters were later shown to provide accurate steady-state data, but the turbine flow meter data should be used in transient phases of operation. Thrust was measured using three load cells in parallel, which also provides the capability to calculate thrust vector alignment. Ignition was demonstrated using a gaseous oxygen/methane spark torch igniter. Test objectives for the RS-18 project are 1) conduct a shakedown of the test stand for LO2/methane lunar ascent engines, 2) obtain vacuum ignition data for the torch and pyrotechnic igniters, and 3) obtain nozzle kinetics data to anchor two-dimensional kinetics codes. All of these objectives were met with the RS-18 data and additional testing data from subsequent LO2/methane test programs in 2009 which included the first simulated-altitude pyrotechnic ignition demonstration of LO2/methane.

Melcher, John C., IV↗
NASA NTRS2009

Sea-Level Flight Demonstration and Altitude Characterization of a LO2/LCH4 Based Accent Propulsion Lander

A 1,500 lbf thrust-class liquid oxygen (LO2)/Liquid Methane (LCH4) rocket engine was developed and tested at both sea-level and simulated altitude conditions. The engine was fabricated by Armadillo Aerospace (AA) in collaboration with NASA Johnson Space Center. Sea level testing was conducted at Armadillo Aerospace facilities at Caddo Mills, TX. Sea-level tests were conducted using both a static horizontal test bed and a vertical take-off and landing (VTOL) test bed capable of lift-off and hover-flight in low atmosphere conditions. The vertical test bed configuration is capable of throttling the engine valves to enable liftoff and hover-flight. Simulated altitude vacuum testing was conducted at NASA Johnson Space Center White Sands Test Facility (WSTF), which is capable of providing altitude simulation greater than 120,000 ft equivalent. The engine tests demonstrated ignition using two different methods, a gas-torch and a pyrotechnic igniter. Both gas torch and pyrotechnic ignition were demonstrated at both sea-level and vacuum conditions. The rocket engine was designed to be configured with three different nozzle configurations, including a dual-bell nozzle geometry. Dual-bell nozzle tests were conducted at WSTF and engine performance data was achieved at both ambient pressure and simulated altitude conditions. Dual-bell nozzle performance data was achieved over a range of altitude conditions from 90,000 ft to 50,000 ft altitude. Thrust and propellant mass flow rates were measured in the tests for specific impulse (Isp) and C* calculations.

Collins, Jacob↗
NASA NTRS2006

Simulator for Testing Spacecraft Separation Devices

A report describes the main features of a system for testing pyrotechnic and mechanical devices used to separate spacecraft and modules of spacecraft during flight. The system includes a spacecraft simulator [also denoted a large mobility base (LMB)] equipped with air thrusters, sensors, and data-acquisition equipment. The spacecraft simulator floats on air bearings over an epoxy-covered concrete floor. This free-flotation arrangement enables simulation of motion in outer space in three degrees of freedom: translation along two orthogonal horizontal axes and rotation about a vertical axis. The system also includes a static stand. In one application, the system was used to test a bolt-retraction system (BRS) intended for separation of the lifting-body and deorbit-propulsion stages of the X- 38 spacecraft. The LMB was connected via the BRS to the static stand, then pyrotechnic devices that actuate the BRS were fired. The separation distance and acceleration were measured. The report cites a document, not yet published at the time of reporting the information for this article, that is said to present additional detailed information.

Johnston, Nick↗
NASA NTRS2003

Swell Sleeves for Testing Explosive Devices

A method of testing explosive and pyrotechnic devices involves exploding the devices inside swell sleeves. Swell sleeves have been used previously for measuring forces. In the present method, they are used to obtain quantitative indications of the energy released in explosions of the devices under test. A swell sleeve is basically a thick-walled, hollow metal cylinder threaded at one end to accept a threaded surface on a device to be tested (see Figure 1). Once the device has been tightly threaded in place in the swell sleeve, the device-and-swell-sleeve assembly is placed in a test fixture, then the device is detonated. After the explosion, the assembly is removed from the test fixture and placed in a coordinate-measuring machine for measurement of the diameter of the swell sleeve as a function of axial position. For each axial position, the original diameter of the sleeve is subtracted from the diameter of the sleeve as swollen by the explosion to obtain the diametral swelling as a function of axial position (see Figure 2). The amount of swelling is taken as a measure of the energy released in the explosion. The amount of swelling can be compared to a standard amount of swelling to determine whether the pyrotechnic device functioned as specified.

Hinkel, Todd J.↗
NASA NTRS2003

Command Interface ASIC - Analog Interface ASIC Chip Set

A command interface application-specific integrated circuit (ASIC) and an analog interface ASIC have been developed as a chip set for remote actuation and monitoring of a collection of switches, which can be used to control generic loads, pyrotechnic devices, and valves in a high-radiation environment. The command interface ASIC (CIA) can be used alone or in combination with the analog interface ASIC (AIA). Designed primarily for incorporation into spacecraft control systems, they are also suitable for use in high-radiation terrestrial environments (e.g., in nuclear power plants and facilities that process radioactive materials). The primary role of the CIA within a spacecraft or other power system is to provide a reconfigurable means of regulating the power bus, actuating all valves, firing all pyrotechnic devices, and controlling the switching of power to all switchable loads. The CIA is a mixed-signal (analog and digital) ASIC that includes an embedded microcontroller with supporting fault-tolerant switch control and monitoring circuitry that is capable of connecting to a redundant set of interintegrated circuit (I(sup 2)C) buses. Commands and telemetry requests are communicated to the CIA. Adherence to the I(sup 2)C bus standard helps to reduce development costs by facilitating the use of previously developed, commercially available components. The AIA is a mixed-signal ASIC that includes the analog circuitry needed to connect the CIA to a custom higher powered version of the I(sup 2)C bus. The higher-powered version is designed to enable operation with bus cables longer than those contemplated in the I(sup 2)C standard. If there are multiple higher-power I(sup 2)C-like buses, then there must an AIA between the CIA and each such bus. The AIA includes two identical interface blocks: one for the side-A I(sup 2)C clock and data buses and the other for the side B buses. All the AIAs on each side are powered from a common power converter module (PCM). Sides A and B of the I(sup 2)C buses are electrically isolated from each other (see figure). They are also isolated from the CIA by use of transformer coupling of signals between the AIA blocks and the CIA.

Ruiz, Baldes↗
NASA NTRS2011

Solid Rocket Booster (SRB) - Evolution and Lessons Learned During the Shuttle Program

The Solid Rocket Booster (SRB) element integrates all the subsystems needed for ascent flight, entry, and recovery of the combined Booster and Motor system. These include the structures, avionics, thrust vector control, pyrotechnic, range safety, deceleration, thermal protection, and retrieval systems. This represents the only human-rated, recoverable and refurbishable solid rocket ever developed and flown. Challenges included subsystem integration, thermal environments and severe loads (including water impact), sometimes resulting in hardware attrition. Several of the subsystems evolved during the program through design changes. These included the thermal protection system, range safety system, parachute/recovery system, and others. Obsolescence issues occasionally required component recertification. Because the system was recovered, the SRB was ideal for data and imagery acquisition, which proved essential for understanding loads and system response. The three main parachutes that lower the SRBs to the ocean are the largest parachutes ever designed, and the SRBs are the largest structures ever to be lowered by parachutes. SRB recovery from the ocean was a unique process and represented a significant operational challenge; requiring personnel, facilities, transportation, and ground support equipment. The SRB element achieved reliability via extensive system testing and checkout, redundancy management, and a thorough postflight assessment process. Assembly and integration of the booster subsystems was a unique process and acceptance testing of reused hardware components was required for each build. Extensive testing was done to assure hardware functionality at each level of stage integration. Because the booster element is recoverable, subsystems were available for inspection and testing postflight, unique to the Shuttle launch vehicle. Problems were noted and corrective actions were implemented as needed. The postflight assessment process was quite detailed and a significant portion of flight operations. The SRBs provided fully redundant critical systems including thrust vector control, mission critical pyrotechnics, avionics, and parachute recovery system. The design intent was to lift off with full redundancy. On occasion, the redundancy management scheme was needed during flight operations. This paper describes some of the design challenges, how the design evolved with time, and key areas where hardware reusability contributed to improved system level understanding.

Kanner, Howard S.↗
NASA NTRS2011

Solid Rocket Booster (SRB) Flight System Integration at Its Best

The Solid Rocket Booster (SRB) element integrates all the subsystems needed for ascent flight, entry, and recovery of the combined Booster and Motor system. These include the structures, avionics, thrust vector control, pyrotechnic, range safety, deceleration, thermal protection, and retrieval systems. This represents the only human-rated, recoverable and refurbishable solid rocket ever developed and flown. Challenges included subsystem integration, thermal environments and severe loads (including water impact), sometimes resulting in hardware attrition. Several of the subsystems evolved during the program through design changes. These included the thermal protection system, range safety system, parachute/recovery system, and others. Because the system was recovered, the SRB was ideal for data and imagery acquisition, which proved essential for understanding loads, environments and system response. The three main parachutes that lower the SRBs to the ocean are the largest parachutes ever designed, and the SRBs are the largest structures ever to be lowered by parachutes. SRB recovery from the ocean was a unique process and represented a significant operational challenge; requiring personnel, facilities, transportation, and ground support equipment. The SRB element achieved reliability via extensive system testing and checkout, redundancy management, and a thorough postflight assessment process. However, the in-flight data and postflight assessment process revealed the hardware was affected much more strongly than originally anticipated. Assembly and integration of the booster subsystems required acceptance testing of reused hardware components for each build. Extensive testing was done to assure hardware functionality at each level of stage integration. Because the booster element is recoverable, subsystems were available for inspection and testing postflight, unique to the Shuttle launch vehicle. Problems were noted and corrective actions were implemented as needed. The postflight assessment process was quite detailed and a significant portion of flight operations. The SRBs provided fully redundant critical systems including thrust vector control, mission critical pyrotechnics, avionics, and parachute recovery system. The design intent was to lift off with full redundancy. On occasion, the redundancy management scheme was needed during flight operations. This paper describes some of the design challenges and technical issues, how the design evolved with time, and key areas where hardware reusability contributed to improved system level understanding.

Wood, T. David↗
NASA NTRS2012

Development and Testing of a Green-Propellant Micro-Hybrid Thruster with Electrostatic Ignition

As early as 1937 German scientists at Peenemunde experimented with highly unstable fuel blends of nitrous oxide (N2O) and ethanol. These early tests mostly resulted in explosions and destroyed rocket engines. More recently several companies have developed experimental nitrous oxide fuel blends (NOFB) with Isp exceeding 300 sec. Although NOFBx has recently been cleared for tests on the International Space Station, this propellant remains highly experimental and has not been cleared for commercial transport by the US DOT. Recent work by Karabeyoglu et al. has raised concerns about the safety risks of mixing hydrocarbons with N2O. Liquid oxidizer/fuel blends are highly explosive and require extreme care in transport and servicing. By adding small amounts of a liquid organic fuel such as alcohol or a hydrocarbon, the odds of an explosive decomposition event are significantly increased.iv The proposed solution mitigates the explosion hazards of NOFB by separating the oxidizer from the hydrocarbon fuel formed as of a small cylindrical section of ABS thermoplastic. As N2O vapor flows across the grain segment, current enters a 1000 VDC high-tension lead in the ABS fuel grain and produces an inductive spark that vaporizes a small amount of the material. The ablated fuel vapor plus residual energy from the spark seed a localized exothermic N2O dissociation that produces sufficient heat to initiate combustion. The process is also effective when gaseous oxygen is used. A low TRL (2-3) prototype demonstrating the feasibility of controlled hydrocarbon-seeding was recently tested at Utah State University.v The unit features a miniature 2.5 cm ABS fuel grain fabricated using a Stratasys Dimension 3-D printer. The 9-N thruster was pulse-fired up to 27 consecutive times on a single ABS grain segment. Ignition was achieved by as little as 12-15 Joules energy input. This value is contrasted with the typical 30-minute pre-heat requirement for the ECAPS LMP-103S ADN-based monopropellant, requiring an energy input of 14,850 Joules for catalytic dissociation. The hydrocarbon-seeded micro-hybrid was also adapted as a non-pyrotechnic ignitor for a 900 N (200-lbf) thrust hybrid motor. The motor was successfully ignited 4 consecutive times with no hardware swaps or propellant additions. The amount of ABS seed material that can be fit into the injector cap is the only limit to the number of available repeat firings. This series of tests marks the first time a hybrid motor was ever ignited by other than a solid-propellant pyrotechnic charge or bi-propellant flame ignitor. Nitrous oxide hybrid motors are typically difficult to ignite and usually require multiple solid-propellant charges to initiate combustion, so this nonpyrotechnic ignition is a significant accomplishment. The controlled hydrocarbon-seeding approach is fundamentally different from all other green propellant solutions offered by the aerospace industry. Although the proposed system is more correctly a hybrid technology; the system retains all the simple features of a monopropellant design. To date no optimization study has been performed to identify the best grain geometry for electrostatic ignition. Fortunately, because the grain segments are fabricated using rapid-prototyping technology, changing the grain geometry is as simple as modifying the 3-D printer CAD-file. Vacuum Isp exceeding 270 seconds has been demonstrated (Ref v), a value significantly higher than those offered by competing green monopropellant options. The propellants of choice, N2O/GOX and ABS are 100% non-toxic, non-explosive, and environmentally benign. Because the inert oxidizer and fuel components are mixed only within the combustion chamber, the system retains the inherent safety of a hybrid rocket and can be piggy-backed as a secondary payload with no overall mission risk increase to the primary payload, an excellent characteristic for secondary launch systems.

Whitmore, Stephen A.↗
NASA NTRS2014

SLM Produced Hermetically Sealed Isolation Valve

Marshall Space Flight Center (MSFC) has developed a valve concept to replace traditional pyrotechnic driven isolation valves. This paper will describe the valve design and development process. The valve design uses a stem/wedge to support a disk inside the valve. That disk hermetically seals the pressurized fluids. A release mechanism holds the stem/wedge and a large spring in place. When required to open, a solenoid is energized and pulls the release mechanism allowing the spring to pull the stem/wedge away from the disk. Now the disk is unsupported and the pressure ruptures the disk allowing flow to the outlet of the valve. This paper will provide details of this design, describe the development testing, and show the results from the valve level tests performed. Also, a trade study is presented to show the advantages of this design to a conventional pyrotechnic based valve.

Richard, James A.↗
NASA NTRS2015

Test and Inspection Requirements for Frangible Joints Used in Human-Rated Spacecraft Applications

The purpose of this document is to define a series of test and inspection requirements for a specific type of frangible joint assembly used in human-rated spacecraft applications. It is intended to be an enhancement to the Commercial Crew Program JSC 62809D Human Rated Spacecraft Pyrotechnic Specification and lists very specific tests and inspections with accompanying tables. This document does not eliminate any of the requirements stated in JSC 62809D. Many of the test and inspections in this document are standard operating procedure for pyrotechnics suppliers. This document provides detail on the individual items to avoid ambiguity. In particular, this document details the test and inspection requirements for offset notch style frangible joints with aluminum machined plates. It is expected that a user would tailor the contents of this document for their particular vehicle application and provide a certifying organization with rationale for such tailoring. Tailoring to the requirements of this document requires approval of the certifying organization (e.g. NASA for a NASA human spaceflight program). The FJ must perform in accordance with design specifications after subjection to all environments it would encounter in its lifecycle. This document does not cover all possible environments that may exist in unique applications and a user may determine that additional tests are necessary. The user must develop written inspection and test procedures with detailed acceptance criteria.

Sirri Oguz↗
NASA NTRS2014

SLM Produced Hermetically Sealed Isolation Valve

Marshall Space Flight Center (MSFC) has developed a valve concept to replace traditional pyrotechnic-driven isolation valves. This paper will describe the valve design and development process. The valve design uses a stem/wedge to support a disk inside the valve. That disk hermetically seals the pressurized fluids. A release mechanism holds the stem/wedge and a large spring in place. When required to open, a solenoid is energized and pulls the release mechanism allowing the spring to pull the stem/wedge away from the disk. Now the disk is unsupported and the pressure ruptures the disk allowing flow to the outlet of the valve. This paper will provide details of this design, describe the development testing, and show the results from the valve level tests performed. Also, a trade study is presented to show the advantages of this design to a conventional pyrotechnic-based valve.

Richard, James↗
NASA NTRS2016

Flat H Frangible Joint Evolution

Space vehicle staging and separation events require pyrotechnic devices. They are single-use mechanisms that cannot be tested, nor can failure-tolerant performance be demonstrated in actual flight articles prior to flight use. This necessitates the implementation of a robust design and test approach coupled with a fully redundant, failure-tolerant explosive mechanism to ensure that the system functions even in the event of a single failure. Historically, NASA has followed the single failure-tolerant (SFT) design philosophy for all human-rated spacecraft, including the Space Shuttle Program. Following the end of this program, aerospace companies proposed building the next generation human-rated vehicles with off-the-shelf, non-redundant, zero-failure-tolerant (ZFT) separation systems. Currently, spacecraft and launch vehicle providers for both the Orion and Commercial Crew Programs (CCPs) plan to deviate from the heritage safety approach and NASA's SFT human rating requirements. Both programs' partners have base-lined ZFT frangible joints for vehicle staging and fairing separation. These joints are commercially available from pyrotechnic vendors. Non-human-rated missions have flown them numerous times. The joints are relatively easy to integrate structurally within the spacecraft. In addition, the separation event is debris free, and the resultant pyro shock is lower than that of other design solutions. It is, however, a serious deficiency to lack failure tolerance. When used for critical applications on human-rated vehicles, a single failure could potentially lead to loss of crew (LOC) or loss of mission (LOM)). The Engineering and Safety & Mission Assurance directorates within the NASA Johnson Space Center took action to address this safety issue by initiating a project to develop a fully redundant, SFT frangible joint design, known as the Flat H. Critical to the ability to retrofit on launch vehicles being developed, the SFT mechanisms must fit within the same three-dimensional envelope as current designs as well as meet structural loads requirements. There is increased mass associated with the redundant design, and the goal is to minimize the weight impact as much as possible. These requirements presented significant challenges, both technically and financially; these challenges will be explored in this paper. Perhaps greater than the technical issues confronted during this design process, were the financial considerations. These were a significant part of the story of this design and development plan. Insufficient financial and labor resources were formidable barriers to completing this project. Nevertheless, JSC personnel successfully conducted several test series at JSC with very useful results. The many lessons learned drove design improvements, performance efficiency, and increased functional reliability. This paper examines the significant technical and financial challenges that these requirements posed to the project team. It discusses the evolution of the SFT frangible joint design, including optimization, testing, and successful partnering of the Johnson Space Center (JSC) engineering and JSC safety organizations, to enhance the flight safety margin for America's next generation of human-rated space vehicles.

Diegelman, Thomas E.↗
NASA NTRS2017

Treatment of Solid Rocket Motors that Complies with Established Protocols to Ensure Planetary Protection

This presentation discusses recent work being conducted by the National Aeronautics and Space Administration (NASA) at Marshall Space Flight Center (MSFC) to evaluate various methods that could be employed to provide for planetary protection of those solar system bodies that are candidates for extraterrestrial life, thus preventing contamination of such bodies. MSFC is presently involved in the development phase of the Europa Lander De-Orbital Stage (DOS) braking motor. In order to prevent bio-contamination of this Jovian satellite, three paths are currently being considered. The first is (1) Bio-Reduction of those microscopic organisms in or on the vehicle (in this case a solid rocket motor (SRM)) that might otherwise be transported during the mission. Possible methods being investigated include heat sterilization, application or incorporation of biocide materials, and irradiation. While each method can be made to work, effects on the SRM's components (propellant, liner, insulation, etc.) could well prove deleterious. A second path would be use of (2) Bio-Barrier material(s). So long as such barrier(s) can maintain their integrity, planetary protection should be afforded. Under the harsh conditions encountered during extended spaceflight (vacuum, temperature extremes, radiation), however, such barrier(s) could well experience a breach. Finally, a third path would be to perform (3) Pyrotechnic Sterilization of the SRM during its end-of-mission phase. Multiple pyrotechnic units would be triggered to ensure activation of such an event and provide for a final sterilization before vehicle impact. In light of Europa's stringent bio-reduction targets, the final and best choice to minimize risk will probably be some combination of the above.

Stefanski, Philip L.↗
NASA NTRS2017

Normally-Closed Zero-Leak Valve with Magnetostrictive Actuator

A non-pyrotechnic, normally-closed, zero-leak valve is a replacement for the pyrovalve used for both in-space and launch vehicle applications. The valve utilizes a magnetostrictive alloy for actuation, rather than pyrotechnic charges. The alloy, such as Terfenol-D, experiences magnetostriction, i.e. a gross elongation, when exposed to a magnetic field. This elongation fractures a parent metal seal, allowing fluid flow through the valve. The required magnetic field is generated by redundant coils that are isolated from the working fluid.

Ramspacher, Daniel J.↗
NASA NTRS2015

Verification and Validation Testing of the Parachute Decelerator System Prior to the First Supersonic Flight Dynamics Test for the Low Density Supersonic Decelerator Program

The Parachute Decelerator System (PDS) is comprised of all components associated with the supersonic parachute and its associated deployment. During the Supersonic Flight Dynamics Test (SFDT), for the Low Density Supersonic Decelerators Program, the PDS was required to deploy the supersonic parachute in a defined fashion. The PDS hardware includes three major subsystems that must function together. The first subsystem is the Parachute Deployment Device (PDD), which acts as a modified pilot deployment system. It is comprised of a pyrotechnic mortar, a Kevlar ballute, a lanyard actuated pyrotechnic inflation aid, and rigging with its associated thermal protection material (TPS). The second subsystem is the supersonic parachute deployment hardware. This includes all of the parachute specific rigging that includes the parachute stowage can and the rigging including TPS and bridle stiffeners for bridle management during deployment. The third subsystem is the Supersonic Parachute itself, which includes the main parachute and deployment bags. This paper summarizes the verification and validation of the deployment process, from the initialization of the PDS system through parachute bag strip that was done prior to the first SFDT.

Gallon, John C.↗
NASA NTRS1966

Spacecraft development

Flight sequence, capsule design, pyrotechnic applications, escape modes, and safety features for Gemini spacecraft

SAFETY DEVICE↗