SEARCH · Engineering Papers
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
We demonstrate an electro-optic microwave modulator with milliWatt control power and a sub-microWatt photonic receiver based on triply-resonant three-wave mixing in high-Q toroidal lithium niobate cavities with whispering-gallery (WG) modes.
A history of learning the operation of Throttles Only Control (TOC) to control an aircraft in real time using engine thrust is shown. The topics include: 1) Past TOC Accidents/Incidents; 2) 1972: DC-10 American Airlines; 3) May 1974: USAF B-52H; 4) April 1975: USAF C-5A; 5) April 1975: USAF C-5A; 6) 1981: USAF B-52G; 7) August 1985: JAL 123 B-747; 8) JAL 123 Survivor Story; 9) JAL 123 Investigation Findings; 10) July 1989: UAL 232 DC-10; 11) UAL 232 DC-10; 12) Eastwind 517 B-737; 13) November 2003: DHL A-300; 14) Historically, TOC has saved lives; 15) Automated Throttles-Only Control; 16) PCA Project; 17) Propulsion-Controlled Aircraft; 18) MD-11 PCA System and Flight Test Envelope; 19) MD-11 Simulation, PCA ILS-Soupled Landing Dispersion; 20) Throttles-Only Pitch and Roll Control Power; 21) PCA in Commercial Fleet; 22) Fall 2005: PCAR Project; 23) PCAR Background - TOC; and 24) PCAR Background - TOC.
An automated mass spectrometer analysis system is disclosed, in which samples are automatically processed in a sample processor and converted into volatilizable samples, or their characteristic volatilizable derivatives. Each volatilizable sample is sequentially volatilized and analyzed in a double focusing mass spectrometer, whose output is in the form of separate ion beams all of which are simultaneously focused in a focal plane. Each ion beam is indicative of a different sample component or different fragments of one or more sample components and the beam intensity is related to the relative abundance of the sample component. The system includes an electro-optical ion detector which automatically and simultaneously converts the ion beams, first into electron beams which in turn produce a related image which is transferred to the target of a vilicon unit. The latter converts the images into electrical signals which are supplied to a data processor, whose output is a list of the components of the analyzed sample and their abundances. The system is under the control of a master control unit, which in addition to monitoring and controlling various power sources, controls the automatic operation of the system under expected and some unexpected conditions and further protects various critical parts of the system from damage due to particularly abnormal conditions.
A system and method for controlling power transferred to an aircraft. The system includes a master antenna and a plurality of slave antennas on the ground. Each slave antenna transmits an uplink signal of a unique phase modulated frequency. The master antenna transmits a master uplink signal. The aircraft receives all the uplink signals and modulates a composite of those signals to produce a downlink beacon that has multiple phase components, each of which corresponds to one of the slave antennas and has a unique frequency. Each of the slave antennas receives the downlink beacon and uses the corresponding phase component to adjust the phase of the slave uplink signal relative to the master uplink signal.
An automated mass spectrometer analysis system is disclosed, in which samples are automatically processed in a sample processor and converted into volatilizable samples, or their characteristic volatilizable derivatives. Each volatizable sample is sequentially volatilized and analyzed in a double focusing mass spectrometer, whose output is in the form of separate ion beams all of which are simultaneously focused in a focal plane. Each ion beam is indicative of a different sample component or different fragments of one or more sample components and the beam intensity is related to the relative abundance of the sample component. The system includes an electro-optical ion detector which automatically and simultaneously converts the ion beams, first into electron beams which in turn produce a related image which is transferred to the target of a vidicon unit. The latter converts the images into electrical signals which are supplied to a data processor, whose output is a list of the components of the analyzed sample and their abundances. The system is under the control of a master control unit, which in addition to monitoring and controlling various power sources, controls the automatic operation of the system under expected and some unexpected conditions and further protects various critical parts of the system from damage due to particularly abnormal conditions.
The Terrestrial Planet Finder - Occulter (TPF-O) mission has two Spacecraft (SC) buses, one for a space telescope and the other for a formation-flying occulter. SC buses typically supply the utilities (support structures, propulsion, attitude control, power, communications, etc) required by the payloads. Unique requirements for the occulter SC bus are to provide the large delta V required for the slewing maneuvers of the occulter, and comunications for formation flying. The TPF-O telescope SC bus shares some key features of the one for the Hubble Space Telescope (HST): both support space telescopes designed to observe in the visible to near infrared range of wavelengths with comparable primary mirror apertures (2.4 m for HST, 2.4 - 4.0 m for TPF-O). However, TPF-O is expected to have a Wide Field Camera (WFC) with a Field of View (FOV) much larger than that of HST. Ths WFC is also expected to provide fine guidance. TPF-O is designed to operate in an orbit around the Sun-Earth Lagrange 2 (SEL2) point. The longer communications range to SEL2 and the large science FOV require higher performance communications than HST. Maintaining a SEL2 orbit requires TPF-O, unlike HST, to have a propulsion system. The velocity required for reachng SEL2 and the limited capabilities of affordable launch vehicles require both TPF-O elements to have compact, low-mass designs. Finally, it is possible that TPF-O may utilize a modular design derived fiom that of HST to allow servicing in the SEL2 orbit.
In 1957 the first earth satellite ushered in the age of space flight. Since that historic event, space exploration has become a major national objective of both the United States and the Soviet Union. These two nations have attempted a total of well over 200 space flight missions. Other nations are also participating in various degrees in what will continue to grow as a cooperative world effort. In the years since 1957, man has successfully flown in earth orbit. He has initiated programs to land on the moon and return. He has made dramatic applications of earth satellites in meteorology, communications, navigation, and geodesy. A host of scientific satellites.continue to advance understanding of the earth's environment, the sun, and the stars. Automated spacecraft are being flown to the moon, deep into interplanetary space, and to the near planets, Mars and Venus. One of the most exciting technological aspects of space exploration has been the development of automated spacecraft. Most of the scientific exploration of space and the useful applications of space flight thus far have been made possible by automated spacecraft. Development of these spacecraft and their many complex subsystems is setting the pace today for many branches of science and technology. Guidance, computer, attitude control, power, telecommunication, instrumentation, and structural subsystems are being subjected to new standards of light weight, high efficiency, extreme accuracy, and unsurpassed reliability and quality. This publication reviews the automated spacecraft which have been developed and flown, or which are under active development in the United States by the National Aeronautics and Space Administration. From the facts and statistics contained herein, certain observations can be made and certain conclusions drawn.
A report describes a rover-type robotic wheeled vehicle recently built for use as a testbed for development of software for future rover-type vehicles. This vehicle is a derivative of the Field Integrated Design and Operations (FIDO) rover, which is a prototype Mars-exploration rover that also serves as a terrestrial testbed. The present vehicle was designed to be nearly functionally identical to the FIDO rover but to be built at much lower cost and to incorporate several improvements to increase utility for development work. Accordingly, considerable effort was made to use commercial off-the-shelf parts and other parts that could be fabricated easily and at low cost. Important features of this vehicle include six-wheel drive and six-wheel steering; onboard computer and power, control, and data-communication electronics having flexibility needed for development of software; significantly increased maximum speed (60 cm/s versus 6 cm/s for the FIDO rover); a rocker-bogey suspension with external differential link, functionally equivalent to that of the FIDO rover; and a hand-held remote controller that can be used to control vehicle motion manually without using the computer (or while waiting for the onboard computer to boot up).
The SDO spacecraft was designed to help us understand the Sun's influence on Earth and Near-Earth space by studying the solar atmosphere on small scales of space and time and in many wavelengths simultaneously. It will perform its operations in a geosynchronous orbit of the earth. This paper will present background on the SDO mission, an overview of the design and development activities associated specifically with the SDO electrical power system (EPS), as well as the major driving requirements behind the mission design. The primary coverage of the paper will be devoted to some of the challenges faced during the design and development phase. This will include the challenges associated with development of a compatible CompactPCI (cPCI) interface within the Power System Electronics (PSE) in order to utilize a "common" processor card, implementation of new solid state power controllers (SSPC) for primary load distribution switching and over current protection in the PSE, and the design approach adopted to meet single fault tolerance requirements for all of the SDO EPS functions.
Boeing and a team from Air Force, NASA, Army, DARPA, MIT, UCLA, and U. of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. The Boeing SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing edge flap on each blade. The eleven-week test program evaluated the forward flight characteristics of the active-flap rotor at speeds up to 155 knots, gathered data to validate state-of-the-art codes for rotor aero-acoustic analysis, and quantified the effects of open and closed loop active flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness of the active flap control on noise and vibration was conclusively demonstrated. Results showed significant reductions up to 6dB in blade-vortex-interaction and in-plane noise, as well as reductions in vibratory hub loads up to 80%. Trailing-edge flap deflections were controlled within 0.1 degrees of the commanded value. The impact of the active flap on control power, rotor smoothing, and performance was also demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than 60 hours of wind-tunnel testing.
This paper describes a novel thermal control system for the Warm Electronics Box (WEB) on board a small lunar surface lander intended to support science activities anywhere on the lunar surface for an extended duration of up to 6 years. Virtually all lander electronics, which collectively dissipate about 60 W in the reference mission, are contained within the WEB. These devices must be maintained below 323 K (with a goal of 303 K) during the nearly 15-earth-day lunar day, when surface temperatures can reach 390K, and above 263 K during the nearly 15-earth-day lunar night, when surface temperatures can reach 100K. Because of the large temperature swing from lunar day-to-night, a novel thermal switching system was required that would be able to provide high conductance from WEB to radiator(s) during the hot lunar day and low (or negligible) conductance during the cold lunar night. The concept that was developed consists of ammonia variable conductance heat pipes (VCHPs) to collect heat from WEB components and a polymer wick propylene loop heat pipe (LHP) to transport the collected heat to the radiator(s). The VCHPs autonomously maximize transport when the WEB is warm and autonomously shut down when the WEB gets cold. The LHP autonomously shuts down when the VCHPs shut down. When the environment transitions from lunar night to day, the VCHPs and LHP autonomously turn back on. Out of 26 analyzed systems, this novel arrangement was able to best achieve the combined goals of zero control power, autonomous operation, long life, low complexity, low T, and landed tilt tolerance.
The next generation space suit requires additional capabilities for controlling and adjusting internal pressure compared to that of historical designs. Next generation suit pressures will range from slight pressure, for astronaut prebreathe comfort, to hyperbaric pressure levels for emergency medical treatment of decompression sickness. Historical regulator designs for space suits have been purely mechanical in nature with the performance aspects typical of such designs such as droop. This paper discusses a new approach to providing the tighter regulation band possible with electronic regulation while mitigating the typical short-comings of electronic control: power consumption, loss of power failure modes, start-up time, and reliability. The paper will also address the bench top prototype development and subsequent performance testing with comparison back to mechanical regulators as a baseline.
Recent tests using the Common Research Model (CRM) at the Langley National Transonic Facility (NTF) and the Ames 11-foot Transonic Wind Tunnel (11' TWT) produced large sets of data that have been used to examine the effects of active damping on transonic tunnel aerodynamic data quality. In particular, large statistically significant sets of repeat data demonstrate that the active damping system had no apparent effect on drag, lift and pitching moment repeatability during warm testing conditions, while simultaneously enabling aerodynamic data to be obtained post stall. A small set of cryogenic (high Reynolds number) repeat data was obtained at the NTF and again showed a negligible effect on data repeatability. However, due to a degradation of control power in the active damping system cryogenically, the ability to obtain test data post-stall was not achieved during cryogenic testing. Additionally, comparisons of data repeatability between NTF and 11-ft TWT CRM data led to further (warm) testing at the NTF which demonstrated that for a modest increase in data sampling time, a 2-3 factor improvement in drag, and pitching moment repeatability was readily achieved not related with the active damping system.
An engineering discipline denoted as hybrid power management (HPM) has emerged from continuing efforts to increase energy efficiency and reliability of hybrid power systems. HPM is oriented toward integration of diverse electric energy-generating, energy-storing, and energy-consuming devices in optimal configurations for both terrestrial and outer-space applications. The basic concepts of HPM are potentially applicable at power levels ranging from nanowatts to megawatts. Potential applications include terrestrial power-generation, terrestrial transportation, biotechnology, and outer-space power systems. Instances of this discipline at prior stages of development were reported (though not explicitly labeled as HPM) in three prior NASA Tech Briefs articles: "Ultracapacitors Store Energy in a Hybrid Electric Vehicle"(LEW-16876), Vol. 24, No. 4 (April 2000), page 63; "Photovoltaic Power Station With Ultracapacitors for Storage" (LEW-17177), Vol. 27, No. 8 (August 2003), page 38; and "Flasher Powered by Photovoltaic Cells and Ultracapacitors" (LEW-17246), Vol. 24, No. 10 (October 2003), page 37. As the titles of the cited articles indicate, the use of ultracapacitors as energy-storage devices lies at the heart of HPM. An ultracapacitor is an electrochemical energy-storage device, but unlike in a conventional rechargeable electrochemical cell or battery, chemical reactions do not take place during operation. Instead, energy is stored electrostatically at an electrode/electrolyte interface. The capacitance per unit volume of an ultracapacitor is much greater than that of a conventional capacitor because its electrodes have much greater surface area per unit volume and the separation between the electrodes is much smaller. Power-control circuits for ultracapacitors can be simpler than those for batteries, for two reasons: (1) Because of the absence of chemical reactions, charge and discharge currents can be greater than those in batteries, limited only by the electrical resistances of conductors; and (2) whereas the charge level of a battery depends on voltage, temperature, age, and load condition, the charge level of an ultracapacitor, like that of a conventional capacitor, depends only on voltage.
Maximum blade loading capability of a coaxial, lift-offset rotor is investigated using a rotorcraft configuration designed in the context of short-haul, medium-size civil and military missions. The aircraft was sized for a 6600-lb payload and a range of 300 nm. The rotor planform and twist were optimized for hover and cruise performance. For the present rotor performance calculations, the collective pitch angle is progressively increased up to and through stall with the shaft angle set to zero. The effects of lift offset on rotor lift, power, controls, and blade airloads and structural loads are examined. The maximum lift capability of the coaxial rotor increases as lift offset increases and extends well beyond the McHugh lift boundary as the lift potential of the advancing blades are fully realized. A parametric study is conducted to examine the differences between the present coaxial rotor and the McHugh rotor in terms of maximum lift capabilities and to identify important design parameters that define the maximum lift capability of the rotor. The effects of lift offset on rotor blade airloads and structural loads are also investigated. Flap bending moment increases substantially as lift offset increases to carry the hub roll moment even at low collective values. The magnitude of flap bending moment is dictated by the lift-offset value (hub roll moment) but is less sensitive to collective and speed.
Re-planning of LTCS TVAC testing and supporting RTA (Receiver Telescope Assembly) Test Plan and Procedure document preparation. The Laser Thermal Control System (LTCS) is designed to maintain the lasers onboard Advanced Topographic Laser Altimeter System (ATLAS) at their operational temperatures. In order to verify the functionality of the LTCS, a thermal balance test of the thermal hardware was performed. During the first cold start of the LTCS, the Loop Heat Pipe (LHP) was unable to control the laser mass simulators temperature. The control heaters were fully on and the loop temperature remained well below the desired setpoint. Thermal analysis of the loop did not show these results. This unpredicted behavior of the LTCS was brought up to a panel of LHP experts. Based on the testing and a review of all the data, there were multiple diagnostic performed in order to narrow down the cause. The prevailing theory is that gravity is causing oscillating flow within the loop, which artificially increased the control power needs. This resulted in a replan of the LTCS test flow and the addition of a GSE heater to allow vertical operation.
This paper addresses the technical aspects of the sample return system for the upcoming Origins, Spectral Interpretation, Resource Identification, and Security-Regolith Explorer (OSIRIS-REx) asteroid sample return mission. The overall mission design and current implementation are presented as an overview to establish a context for the technical description of the reentry and landing segment of the mission.The prime objective of the OSIRIS-REx mission is to sample a primitive, carbonaceous asteroid and to return that sample to Earth in pristine condition for detailed laboratory analysis. Targeting the near-Earth asteroid Bennu, the mission launches in September 2016 with an Earth reentry date of September 24, 2023.OSIRIS-REx will thoroughly characterize asteroid Bennu providing knowledge of the nature of near-Earth asteroids that is fundamental to understanding planet formation and the origin of life. The return to Earth of pristine samples with known geologic context will enable precise analyses that cannot be duplicated by spacecraft-based instruments, revolutionizing our understanding of the early Solar System. Bennu is both the most accessible carbonaceous asteroid and one of the most potentially Earth-hazardous asteroids known. Study of Bennu addresses multiple NASA objectives to understand the origin of the Solar System and the origin of life and will provide a greater understanding of both the hazards and resources in near-Earth space, serving as a precursor to future human missions to asteroids.This paper focuses on the technical aspects of the Sample Return Capsule (SRC) design and concept of operations, including trajectory design and reentry retrieval. Highlights of the mission are included below.The OSIRIS-REx spacecraft provides the essential functions for an asteroid characterization and sample return mission: attitude control propulsion power thermal control telecommunications command and data handling structural support to ensure successful rendezvous with Bennu characterization of Bennus properties delivery of the sampler to the surface, and return of the spacecraft to the vicinity of the Earth sample collection, performed by the Touch-and-Go Sample Acquisition Mechanism (TAGSAM), to acquire a regolith sample from the surface Earth re-entry and SRC recovery. Following sample collection, OSIRIS-REx drifts away from Bennu until the Asteroid Departure Maneuver is commanded on March 4, 2021, sending OSIRIS-REx on a ballistic return cruise to Earth. No additional large deterministic maneuvers are required to return the SRC to Earth. During the cruise, tracking and trajectory correction maneuvers (TCMs) are performed as necessary to precisely target the entry corridor. As OSIRIS-REx approaches Earth, the reentry plans are reviewed starting about a year before arrival, and preparations begin. The spacecraft is targeted away from the Earth until 7 days before entry. The final two trajectory correction maneuvers bring the spacecraft on target toward the Utah Test and Training Range (UTTR), with sufficient time for contingency resolution. The SRC releases 4 hours prior to atmospheric entry interface and, using the Stardust capsule heritage design, employs a traditional drogue and main parachute descent system for a soft touchdown.
We present the concept of using a variant of a Space Exploration Technologies Corporation (SpaceX) Dragon space capsule as a low-cost, large-capacity, near-term, Mars lander (dubbed "Red Dragon") for scientific and human precursor missions. SpaceX initially designed the Dragon capsule for flight near Earth, and Dragon has successfully flown many times to low-Earth orbit (LEO) and successfully returned the Dragon spacecraft to Earth. Here we present capsule hardware modifications that are required to enable flight to Mars and operations on the martian surface. We discuss the use of the Dragon system to support NASA Discovery class missions to Mars and focus in particular on Dragon's applications for drilling missions. We find that a Red Dragon platform is well suited for missions capable of drilling deeper on Mars (at least 2 m) than has been accomplished to date due to its ability to land in a powered controlled mode, accommodate a long drill string, and provide payload space for sample processing and analysis. We show that a Red Dragon drill lander could conduct surface missions at three possible targets including the ice-cemented ground at the Phoenix landing site (68degN), the subsurface ice discovered near the Viking 2 (49degN) site by fresh impact craters, and the dark sedimentary subsurface material at the Curiosity site (4.5degS).