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NASA NTRS2005

Beyond the Prototype: The Design Evolution of a Deployed AI System

Our previous experiences with deployed intelligent control agents for NASA advanced life support systems (Schreckenghost et al., 2002) inspired us to develop the Distributed Collaboration and Interaction (DCI) system to help humans and mostly-autonomous software agents work together. We discovered many unaddressed needs for human interaction with control agents that operate continuously over months to years to monitor and perform process control for regenerative life support systems. These systems recover usable water or air from the waste products created by biological systems over time. Through the DCI project, we have addressed the needs for interaction between humans and autonomous control systems. This paper describes the software engineering aspects of our experiences, first, in designing and developing a prototype of the DCI system, and later, in adjusting the implementation to make the leap from initial prototype to a system ready to be applied under varying circumstances to meet different needs. The details in the paper focus primarily on the second stage of development. The DCI system is currently deployed, operating 24/7, to assist humans to interact with an advanced Water Recovery System (WRS). In addition, we have demonstrated an application of DCI to support ground personnel in mission support roles.

Martin, C.↗

NASA NTRS2006

Mechanical Technology Development on A 35-m Deployable Radar Antenna for Monitoring Hurricanes

The NEXRAD in Space project develops a novel instrument concept and the associated antenna technologies for a 35-GHz Doppler radar to monitor hurricanes, cyclones, and severe storms from a geostationary orbit. Mechanical challenges of this concept include a 35-m diameter lightweight in space deployable spherical reflector and a feeder scanning mechanism. The feasibility of using shape memory polymer material to develop the large deployable reflector has been investigated by this study. A spiral scanning mechanism concept has been developed and demonstrated by an engineering model.

radar antennas↗

NASA NTRS2004

Beyond the Prototype: The Design Evolution of a Deployed AI System

Martin, Cheryl↗

NASA NTRS1968

Flight Test of a 40-Foot Nominal-Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 1.91 and a Dynamic Pressure of 11.6 Pounds per Square Foot

A 40-foot (12.2 meter) nominal-diameter disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator Program (SPED-I). The test parachute was ejected by a deployment mortar from an instrumented payload at an altitude of 140,000 feet (42.5 kilometers). The payload was at a Mach number of 1.91 and the dynamic pressure was 11.6 pounds per square foot (555 newtons per square meter) at the time the parachute deployment mortar was fired. The parachute reached suspension line stretch in 0.43 second with a resultant snatch force loading of 1990 pounds (8850 newtons). The maximum parachute opening load of 6500 pounds (28,910 newtons) came 0.61 second later at a total elapsed time from mortar firing of 1.04 seconds. The first full inflation occurred at 1.12 seconds and stable inflation was achieved at approximately 1.60 seconds. The parachute had an average axial-force coefficient of 0.53 during the deceleration period. During the steady-state descent portion of the flight test, the average effective drag coefficient was also 0.53 and pitch-yaw oscillations of the canopy averaged less than 10 degrees in the altitude region above 100,000 feet (30.5 meters).

EXPERIMENTS↗

NASA NTRS1967

Flight Tests of a 40-Foot Nominal Diameter Modified Ringsail Parachute Deployed at Mach 1.64 and Dynamic Pressure of 9.1 Pounds Per Square Foot

A ringsail parachute, which had a nominal diameter of 40 feet (12.2 meters) and reference area of 1256 square feet (117 m(exp 2)) and was modified to provide a total geometric porosity of 15 percent of the reference area, was flight tested as part of the rocket launch portion of the NASA Planetary Entry Parachute Program. The payload for the flight test was an instrumented capsule from which the test parachute was ejected by a deployment mortar when the system was at a Mach number of 1.64 and a dynamic pressure of 9.1 pounds per square foot (43.6 newtons per m(exp 2)). The parachute deployed to suspension line stretch in 0.45 second with a resulting snatch force of 1620 pounds (7200 newtons). Canopy inflation began 0.07 second later and the parachute projected area increased slowly to a maximum of 20 percent of that expected for full inflation. During this test, the suspension lines twisted, primarily because the partially inflated canopy could not restrict the twisting to the attachment bridle and risers. This twisting of the suspension lines hampered canopy inflation at a time when velocity and dynamic-pressure conditions were more favorable.

TESTS↗

NASA NTRS1967

Flight Test of a 30-Foot Nominal-Diameter Disk-Gap-Band Parachute Deployed at Mach 1.56 and Dynamic Pressure of 11.4 Pounds per Square Foot

A 30-foot (9.1 meter) nominal-diameter disk-gap-band parachute (reference area 707 sq ft (65.7 m(exp 2)) was flight tested with a 200-pound (90.7 kg) instrumented payload as part of the NASA Planetary Entry Parachute Program. A deployment mortar ejected the test parachute when the payload was at a Mach number of 1.56 and a dynamic pressure of 11.4 lb/sq ft (546 newtons per m 2 ) at an altitude of 127,500 feet (38.86 km). The parachute reached suspension line stretch in 0.37 second resulting in a snatch force loading of 1270 pounds (5650 N). Canopy inflation began 0.10 second after line stretch. A delay in the opening process occurred and was apparently due to a momentary interference of the glass-fiber shroud used in packing the parachute bag in the mortar. Continuous canopy inflation began 0.73 second after initiation of deployment and 0.21 second later full inflation was attained for a total elapsed time from mortar fire of 0.94 second. The maximum opening load of 3915 pounds (17,400 newtons) occurred at the time the canopy was first fully opened. The parachute exhibited an average drag coefficient of 0.52 during the deceleration period and pitch-yaw oscillations of the canopy were less than 5 degrees. During the steady-state descent portion of the test period, the average effective drag coefficient was about 0.47 (based on vertical descent velocity and total system weight).

EXPERIMENTS↗

NASA NTRS1967

Flight Test of 31.2 Diameter Modified Ringsail Parachute Deployed at Mach 1.39, Dynamic Pressure 11 Pounds per Square Foot

A 31.2-foot (9.51 meter) nominal diameter (reference area 764 ft(exp 2) (71.0 m(exp 2)) ringsail parachute modified to provide 15-percent geometric porosity was flight tested while attached to a 201-pound mass (91.2 kilogram) instrumented payload as part of the rocket launch portion of the NASA Planetary Entry Parachute Program (PEPP). The parachute deployment was initiated by the firing of a mortar at a Mach number of 1.39 and a dynamic pressure of 11.0 lb/ft(exp 2) (527 newtons/m(exp 2)) at an altitude of 122,500 feet (37.3 kilometers). The parachute deployed to suspension-line stretch (snatch force) in 0.35 second, and 0.12 second later the drag force increase associated with parachute inflation began. The parachute inflated in 0.24 second to the full-open condition for a total elapsed opening time of 0.71 second. The maximum opening load of 3970 pounds (17,700 newtons) came at the time the parachute was just fully opened. During the deceleration period, the parachute exhibited an average drag coefficient of 0.52 and oscillations of the parachute canopy were less than 5 degrees. During the steady-state terminal descent portion of the test period, the average effective drag coefficient (based on vertical descent velocity) was 0.52.

TESTS↗

NASA NTRS1971

High Altitude Flight Test of a Reefed 12.2 Meter Diameter Disk-Gap-Band Parachute with Deployment at Mach Number of 2.58

A reefed 12.2-meter nominal-diameter (40-ft) disk-gap-band parachute was flight tested as part of the NASA Supersonic High Altitude Parachute Experiment (SHAPE) program. A three-stage rocket was used to drive the instrumented payload to an altitude of 43.6 km (143,000 ft), a Mach number of 2.58, and a dynamic pressure of 972 N/m(exp 2) (20.3 lb/ft(exp 2)) where the parachute was deployed by means of a mortar. The parachute deployed satisfactorily and reached a partially inflated condition characterized by irregular variations in parachute projected area. A full, stable reefed inflation was achieved when the system had decelerated to a Mach number of about 1.5. The steady, reefed projected area was 49 percent of the steady, unreefed area and the average drag coefficient was 0.30. Disreefing occurred at a Mach number of 0.99 and a dynamic pressure of 81 N/m(exp 2) (1.7 lb/ft(exp 2)). The parachute maintained a steady inflated shape for the remainder of the deceleration portion of the flight and throughout descent. During descent, the average effective drag coefficient was 0.57. There was little, if any, coning motion, and the amplitude of planar oscillations was generally less than 10 degrees. The film also shows a wind tunnel test of a 1.7-meter-diameter parachute inflating at Mach number 2.0.

EXPERIMENTS↗

NASA NTRS1969

Flight Tests Results from Supersonic Deployment of an 18-Foot Diameter (5.49 meter) Towed Ballute Decelerator

A ram-air-inflated, towed ballute decelerator having a maximum frontal diameter of 18 feet (5.49 meters) was deployed during free flight at a Mach number of 3.15 and a dynamic pressure of 38.5 lb/ft(exp 2) (1843.4 newtons/m(exp 2)). Deployment and extraction of the test ballute were normal but inflation stopped about 1 second after mortar firing and produced an average plateau drag force of 1500 pounds (6.7 kN) for about 1 second. Approximately 30 percent of expected total frontal area was obtained.

TESTS↗

NASA NTRS2007

Development of a Deployable Nonmetallic Boom for Reconfigurable Systems of Small Modular Spacecraft

Launch vehicle payload capacity and the launch environment represent two of the most operationally limiting constraints on space system mass, volume, and configuration. Large-scale space science and power platforms as well as transit vehicles have been proposed that greatly exceed single-launch capabilities. Reconfigurable systems launched as multiple small modular spacecraft with the ability to rendezvous, approach, mate, and conduct coordinated operations have the potential to make these designs feasible. A key characteristic of these proposed systems is their ability to assemble into desired geometric (spatial) configurations. While flexible and sparse formations may be realized by groups of spacecraft flying in close proximity, flyers physically connected by active structural elements could continuously exchange power, fluids, and heat (via fluids). Configurations of small modular spacecraft temporarily linked together could be sustained as long as needed with minimal propellant use and reconfigured as often as needed over extended missions with changing requirements. For example, these vehicles could operate in extremely compact configurations during boost phases of a mission and then redeploy to generate power or communicate while coasting and upon reaching orbit. In 2005, NASA funded Phase 1 of a program called Modular Reconfigurable High-Energy Technology Demonstrator Assembly Testbed (MRHE) to investigate reconfigurable systems of small spacecraft. The MRHE team was led by NASA's Marshall Space Flight Center and included Lockheed Martin's Advanced Technology Center (ATC) in Palo Alto and its subcontractor, ATK. One of the goals of Phase 1 was to develop an MRHE concept demonstration in a relevant 1-g environment to highlight a number of requisite technologies. In Phase 1 of the MRHE program, Lockheed Martin devised and conducted an automated space system assembly demonstration featuring multipurpose free-floating robots representing Spacecraft in the newly built Controls and Automation Laboratory (CAL) at the ATC. The CAL lab features a 12' x 24' granite air-bearing table and an overhead simulated starfield. Among the technologies needed for the concept demo were mating interfaces allowing the spacecraft to dock and deployable structures allowing for adjustable separation between spacecraft after a rigid connection had been established. The decision to use a nonmetallic deployable boom for this purpose was driven by the MRHE concept demo requirements reproduced in Table 1.

Rehnmark, Fredrik↗

NASA NTRS2006

Surface Control of Cold Hibernated Elastic Memory Self-Deployable Structure

A new class of simple, reliable, lightweight, low packaging volume and cost, self-deployable structures has been developed for use in space and commercial applications. This technology called 'cold hibernated elastic memory' (CHEM) utilizes shape memory polymers (SMP)in open cellular (foam) structure or sandwich structures made of shape memory polymer foam cores and polymeric composite skins. Some of many potential CHEM space applications require a high precision deployment and surface accuracy during operation. However, a CHEM structure could be slightly distorted by the thermo-mechanical processing as well as by thermal space environment Therefore, the sensor system is desirable to monitor and correct the potential surface imperfection. During these studies, the surface control of CHEM smart structures was demonstrated using a Macro-Fiber Composite (MFC) actuator developed by the NASA LaRC and US Army ARL. The test results indicate that the MFC actuator performed well before and after processing cycles. It reduced some residue compressive strain that in turn corrected very small shape distortion after each processing cycle. The integrated precision strain gages were detecting only a small flat shape imperfection indicating a good recoverability of original shape of the CHEM test structure.

glass transition temperature Tg↗

NASA NTRS2003

Deployable radiator with flexible line loop

Radiator assembly (10) for use on a spacecraft (12) is provided including at least one radiator panel assembly (26) repeatably movable between a panel stowed position (28) and a panel deployed position (36), at least two flexible lines (40) in fluid communication with the at least one radiator panel assembly (26) and repeatably movable between a stowage loop (42) and a flattened deployed loop (44).

Keeler, Bryan V.↗

NASA NTRS2006

Deployable antenna

A deployable antenna and method for using wherein the deployable antenna comprises a collapsible membrane having at least one radiating element for transmitting electromagnetic waves, receiving electromagnetic waves, or both.

Fink, Patrick W.↗

NASA NTRS2008

Problems With Deployment of Multi-Domained, Multi-Homed Mobile Networks

This document describes numerous problems associated with deployment of multi-homed mobile platforms consisting of multiple networks and traversing large geographical areas. The purpose of this document is to provide insight to real-world deployment issues and provide information to groups that are addressing many issues related to multi-homing, policy-base routing, route optimization and mobile security - particularly those groups within the Internet Engineering Task Force.

Ivancic, William D.↗

NASA NTRS2009

Design and Performance of a Spectrometer for Deployment on MISSE 7

A spectrometer for reflectance and transmission measurements of samples exposed to the space environment has been developed for deployment on the Materials on the International Space Station Experiment (MISSE) 7. The instrument incorporates a miniature commercial fiber optic coupled spectrometer with a computer control system for detector operation, sample motion and illumination. A set of three spectrometers were recently integrated on the MISSE7 platform with launch and deployment on the International Space Station scheduled for summer of this year. The instrument is one of many active experiments on the platform. The performance of the instrument prior to launch will be discussed. Data from samples measured in the laboratory will be compared to those from the instrument prior to launch. These comparisons will illustrate the capabilities of the current design. The space environment challenges many materials. When in operation on the MISSE 7 platform, the new spectrometer will provide real time data on the how the space environment affects the optical properties of thermal control paints and optical coatings. Data obtained from comparison of pre and post flight measurements on hundreds of samples exposed on previous MISSE platforms have been reported at these meetings. With the new spectrometer and the ability to correlate measured changes with time on orbit and the occurrence of both natural events and human activities, a better understanding of the processes responsible for degradation of materials in space will be possible.

Pippin, Gary↗

NASA NTRS2009

Challenges with Deploying and Integrating Environmental Control and Life Support Functions in a Lunar Architecture with High Degrees of Mobility

Visions of lunar outposts often depict a collection of fixed elements such as pressurized habitats, in and around which human inhabitants spend the large majority of their surface stay time. In such an outpost, an efficient deployment of environmental control and life support equipment can be achieved by centralizing certain functions within one or a minimum number of habitable elements and relying on the exchange of gases and liquids between elements via atmosphere ventilation and plumbed interfaces. However, a rigidly fixed outpost can constrain the degree to which the total lunar landscape can be explored. The capability to enable widespread access across the landscape makes a lunar architecture with a high degree of surface mobility attractive. Such mobility presents unique challenges to the efficient deployment of environmental control and life support functions in multiple elements that may for long periods of time be operated independently. This paper describes some of those anticipated challenges.

Bagdigian, Robert M.↗

NASA NTRS2010

A Novel Approach for a Low-Cost Deployable Antenna

The Naval Research Laboratory (NRL) has designed, built, and fully qualified a low cost, low Passive Intermodulation (PIM) 12-foot (3.66-m) diameter deployable ultra high frequency (UHF) antenna for the Tacsat-4 program. The design utilized novel approaches in reflector material and capacitive coupling techniques. This paper discusses major design trades, unique design characteristics, and lessons learned from the development of the Tacsat 4 deployable antenna. This antenna development was sponsored by the Office of Naval Research.

Amend, Chris↗

NASA NTRS2010

Xatcobeo: Small Mechanisms for CubeSat Satellites - Antenna and Solar Array Deployment

The Xatcobeo project, which includes the mechanisms dealt with here, is principally a university project to design and construct a CubeSat 1U-type satellite. This work describes the design and operational features of the system for antenna storage and deployment, and the design and simulations of the solar array deployment system. It explains the various problems faced and solutions adopted, with a view to providing valid data for any other applications that could find them useful, be they of a similar nature or not.

EncinasPlaza, Jose Miguel↗