Project UM-HAUL: A self-unloading reusable lunar lander

NASA Technical Reports Server (NTRS)

1991-01-01

The establishment of a lunar base is technologically and financially challenging. Given the necessary resources and political support, it can be done. In addition to the geopolitical obstacles, however, there are logistical problems involved in establishing such bases that can only be overcome with the acquisition of a significant transportation and communications network in the Earth-Moon spatial region. Considering the significant number of payloads that will be required in this process, the mass-specific cost of launching these payloads, and the added risk and cost of human presence in space, it is clearly desirable to automate major parts of such an operation. One very costly and time-consuming factor in this picture is the delivery of payloads to the Moon. Foreseeable payloads would include atmospheric modules, inflatable habitat kits, energy and oxygen plant elements, ground vehicles, laboratory modules, crew supplies, etc. The duration of high-risk human presence on the Moon could be greatly reduced if all such payloads were delivered to the prospective base site in advance of crew arrival. In this view, the idea of a 'Self-Unloading Reusable Lunar Lander' (SURLL) arises naturally. The general scenario depicts the lander being brought to low lunar orbit (LLO) from Earth atop a generic Orbital Transfer Vehicle (OTV). From LLO, the lander shuttles payloads down to the lunar surface, where, by means of some resident, detachable unloading device, it deploys the payloads and returns to orbit. The general goal is for the system to perform with maximum payload capability, automation, and reliability, while also minimizing environmental hazards, servicing needs, and mission costs. Our response to this demand is UM-HAUL, or the UnManned Heavy pAyload Unloader and Lander. The complete study includes a system description, along with a preliminary cost analysis and a design status assessment.

Project UM-HAUL: A self-unloading reusable lunar lander

NASA Astrophysics Data System (ADS)

The establishment of a lunar base is technologically and financially challenging. Given the necessary resources and political support, it can be done. In addition to the geopolitical obstacles, however, there are logistical problems involved in establishing such bases that can only be overcome with the acquisition of a significant transportation and communications network in the Earth-Moon spatial region. Considering the significant number of payloads that will be required in this process, the mass-specific cost of launching these payloads, and the added risk and cost of human presence in space, it is clearly desirable to automate major parts of such an operation. One very costly and time-consuming factor in this picture is the delivery of payloads to the Moon. Foreseeable payloads would include atmospheric modules, inflatable habitat kits, energy and oxygen plant elements, ground vehicles, laboratory modules, crew supplies, etc. The duration of high-risk human presence on the Moon could be greatly reduced if all such payloads were delivered to the prospective base site in advance of crew arrival. In this view, the idea of a 'Self-Unloading Reusable Lunar Lander' (SURLL) arises naturally. The general scenario depicts the lander being brought to low lunar orbit (LLO) from Earth atop a generic Orbital Transfer Vehicle (OTV). From LLO, the lander shuttles payloads down to the lunar surface, where, by means of some resident, detachable unloading device, it deploys the payloads and returns to orbit. The general goal is for the system to perform with maximum payload capability, automation, and reliability, while also minimizing environmental hazards, servicing needs, and mission costs. Our response to this demand is UM-HAUL, or the UnManned Heavy pAyload Unloader and Lander. The complete study includes a system description, along with a preliminary cost analysis and a design status assessment.

Design feasibility via ascent optimality for next-generation spacecraft

NASA Astrophysics Data System (ADS)

Miele, A.; Mancuso, S.

This paper deals with the optimization of the ascent trajectories for single-stage-sub-orbit (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem is studied for different values of the engine specific impulse and spacecraft structural factor. The main conclusions are that: feasibility of SSSO spacecraft is guaranteed for all the parameter combinations considered; feasibility of SSTO spacecraft depends strongly on the parameter combination chosen; not only feasibility of TSTO spacecraft is guaranteed for all the parameter combinations considered, but the TSTO payload is several times the SSTO payload. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, aerodynamic improvements do not yield significant improvements in payload. For SSSO, SSTO, and TSTO spacecraft, simple engineering approximations are developed connecting the maximum payload weight to the engine specific impulse and spacecraft structural factor. With reference to the specific impulse/structural factor domain, these engineering approximations lead to the construction of zero-payload lines separating the feasibility region (positive payload) from the unfeasibility region (negative payload).

Upper stages utilizing electric propulsion

NASA Technical Reports Server (NTRS)

Byers, D. C.

1980-01-01

The payload characteristics of geocentric missions which utilize electron bombardment ion thruster systems are discussed. A baseline LEO to GEO orbit transfer mission was selected to describe the payload capabilities. The impacts on payloads of both mission parameters and electric propulsion technology options were evaluated. The characteristics of the electric propulsion thrust system and the power requirements were specified in order to predict payload mass. This was completed by utilizing a previously developed methodology which provides a detailed thrust system description after the final mass on orbit, the thrusting time, and the specific impulse are specified. The impact on payloads of total mass in LEO, thrusting time, propellant type, specific impulse, and power source characteristics was evaluated.

STS propellant densification feasibility study data book

NASA Technical Reports Server (NTRS)

Fazah, M. M.

1994-01-01

The feasibility of using densification or subcooling with respect to standard temperature propellants on the Space Transportation System (STS) in order to achieve a payload gain is discussed in this report. The objective is to determine the magnitude of the payload gain and to identify any system impacts to the space shuttle on either flight systems or ground systems. Results show that a payload benefit can be obtained by subcooling the liquid hydrogen (LH2) from a nominal temperature of 36.4 R to 28.5 R and by subcooling the liquid oxygen (LO2) from a nominal temperature of 164 R to either 132.1 R or 141.4 R. When the propellants are subcooled to 28.5 R and 132.1 R for the LH2 and LO2, respectively, a maximum payload gain of 7,324 lb can be achieved, and when the propellants are subcooled to 28.5 R and 141.5 R for the LH2 and LO2, respectively, a maximum payload gain of 6,841 lb can be achieved. If the LH2 is subcooled to 28.5 R while the LH2 and LO2 remains at the nominal conditions, a maximum payload gain of 1,303 lb can be achieved.

Evaluation of propellant tank insulation concepts for low-thrust chemical propulsion systems

NASA Technical Reports Server (NTRS)

Kramer, T.; Brogren, E.; Seigel, B.

1984-01-01

An analytical evaluation of cryogenic propellant tank insulations for liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) was conducted. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three Shuttle-launched LTPS designed for Shuttle bay packaged payload densities of 56 kg/cu m, 40 kg/cu m and 24 kg/cu m. Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because Orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. It was noted this cost could be reduced significantly without increasing program risk.

Electric propulsion for geostationary orbit insertion

NASA Technical Reports Server (NTRS)

Oleson, Steven R.; Curran, Francis M.; Myers, Roger M.

1995-01-01

Solar electric propulsion (SEP) technology is already being used for geostationary satellite stationkeeping to increase payload mass. By using this same technology to perform part of the orbit transfer additional increases in payload mass can be achieved. Advanced chemical and N2H4 arcjet systems are used to increase the payload mass by performing stationkeeping and part of the orbit transfer. Four mission options are analyzed which show the impact of either sharing the orbit transfer between chemical and SEP systems or having either complete the transfer alone. Results show that for an Atlas 2AS payload increases in net mass (geostationary satellite mass less wet propulsion system mass) of up to 100 kg can be achieved using advanced chemical for the transfer and advanced N2H4 arcjets for stationkeeping. An additional 100 kg can be added using advanced N2H4 arcjets for part of a 40 day orbit transfer.

The use of artificial intelligence techniques to improve the multiple payload integration process

NASA Technical Reports Server (NTRS)

Cutts, Dannie E.; Widgren, Brian K.

1992-01-01

A maximum return of science and products with a minimum expenditure of time and resources is a major goal of mission payload integration. A critical component then, in successful mission payload integration is the acquisition and analysis of experiment requirements from the principal investigator and payload element developer teams. One effort to use artificial intelligence techniques to improve the acquisition and analysis of experiment requirements within the payload integration process is described.

Overview of the NASA Entry, Descent and Landing Systems Analysis Exploration Feed-Forward Study

NASA Technical Reports Server (NTRS)

DwyerCianciolo, Alicia M.; Zang, Thomas A.; Sostaric, Ronald R.; McGuire, M. Kathy

2011-01-01

Technology required to land large payloads (20 to 50 mt) on Mars remains elusive. In an effort to identify the most viable investment path, NASA and others have been studying various concepts. One such study, the Entry, Descent and Landing Systems Analysis (EDLSA) Study [1] identified three potential options: the rigid aeroshell, the inflatable aeroshell and supersonic retropropulsion (SRP). In an effort to drive out additional levels of design detail, a smaller demonstrator, or exploration feed-forward (EFF), robotic mission was devised that utilized two of the three (inflatable aeroshell and SRP) high potential technologies in a configuration to demonstrate landing a two to four metric ton payload on Mars. This paper presents and overview of the maximum landed mass, inflatable aeroshell controllability and sensor suite capability assessments of the selected technologies and recommends specific technology areas for additional work.

Optimizing the Space Transportation System

DTIC Science & Technology

1982-12-01

the entire gamut of inclinations and altitudes. The Shuttle payload mass and the total Station, OTV and Shuttle propellant mass required in orbit are...military satellites across the entire gamut of inclinations and altitudes. The total Station, OTV, Shuttle propellant and Shuttle payload mass required in

Oxygen Generator System Mars In-Situ Propellant Production Precursor Flight

NASA Technical Reports Server (NTRS)

Sridhar, K. R.; Gottmann, M.; Baird, R. S.

1999-01-01

The 2001 Lander to Mars will carry the first ever ISRU payload to Mars. This payload, the Mars In-situ Propellant production Precursor (MIP), will demonstrate a variety of technologies that will be required for future ISRU Mars indigenous material processing plant designs. One of those technologies is that of extracting oxygen from the predominantly carbon dioxide atmosphere of Mars, a prerequisite for future sample return and human missions to Mars. The Oxygen Generator Subsystem (OGS) portion of the MIP will demonstrate this and is the focus of this paper. The primary objective of the OGS is to demonstrate the production of oxygen from Mars atmospheric gases. Secondary objectives are to measure the performance and reliability of oxygen generation hardware in actual mission environments over an extended time. Major constraints on the OGS design came from several sources. The Lander provides power to the system from solar power that is harnessed by photovoltaic arrays. This limited OGS to daytime only operations (six to eight hours) and a maximum power of 15W. The reliance on solar power necessitated thermal cycling of the OGS between Mars ambient and OGS operating temperatures. The Lander also limited the total mass of the MIP payload to 7.5 kg with a correspondingly small volume, and the OGS was one of six experiments in the MIP. Mass and volume were to be minimized. Another constraint was cost. Mission funding, as always, was tight. Cost was to be minimized. In short the OGS design had to be low power (<15 Watts), low mass (1 kg), low volume, low cost, and be capable of cyclical operations for an extended stay on Mars. After extensive research, a zirconia based solid oxide electrolyzer design was selected.

Oxygen Generator System Mars In-Situ Propellant Production Precursor Flight

NASA Technical Reports Server (NTRS)

Sridhar, K. R.; Gottmann, M.; Baird, R. S.

1999-01-01

The 2001 Lander to Mars will carry the first ever In situ Resource Utilization (ISRU) payload to Mars. This payload, the Mars In-situ Propellant production Precursor (MIP), will demonstrate a variety of technologies that will be required for future ISRU Mars indigenous material processing plant designs. One of those technologies is that of extracting oxygen from the predominantly carbon dioxide atmosphere of Mars, a prerequisite for future sample return and human missions to Mars. The Oxygen Generator Subsystem (OGS) portion of the MIP will demonstrate this and is the focus of this paper. The primary objective of the OGS is to demonstrate the production of oxygen from Mars atmospheric gases. Secondary objectives are to measure the performance and reliability of oxygen generation hardware in actual mission environments over an extended time. Major constraints on the OGS design came from several sources. The Lander provides power to the system from solar power that is harnessed by photovoltaic arrays. This limited OGS to daytime only operations (six to eight hours) and a maximum power of 15W. The reliance on solar power necessitated thermal cycling of the OGS between Mars ambient and OGS operating temperatures. The Lander also limited the total mass of the MIP payload to 7.5 kg with a correspondingly small volume, and the OGS was one of six experiments in the MIP Mass and volume were to be minimized. Another constraint was cost. Mission funding, as always, was tight. Cost was to be minimized. In short the OGS design had to be low power (<15 Watts), low mass (1 kg), low volume, low cost, and be capable of cyclical operations for an extended stay on Mars. After extensive research, a zirconia based solid oxide electrolyzer design was selected.

Quadrotor Control in the Presence of Unknown Mass Properties

NASA Astrophysics Data System (ADS)

Duivenvoorden, Rikky Ricardo Petrus Rufino

Quadrotor UAVs are popular due to their mechanical simplicity, as well as their capability to hover and vertically take-off and land. As applications diversify, quadrotors are increasingly required to operate under unknown mass properties, for example as a multirole sensor platform or for package delivery operations. The work presented here consists of the derivation of a generalized quadrotor dynamic model without the typical simplifying assumptions on the first and second moments of mass. The maximum payload capacity of a quadrotor in hover, and the observability of the unknown mass properties are discussed. A brief introduction of L1 adaptive control is provided, and three different L 1 adaptive controllers were designed for the Parrot AR.Drone quadrotor. Their tracking and disturbance rejection performance was compared to the baseline nonlinear controller in experiments. Finally, the results of the combination of L1 adaptive control with iterative learning control are presented, showing high performance trajectory tracking under uncertainty.

Evaluation of propellent tank insulation concepts for low-thrust chemical propulsion systems: Executive summary

NASA Technical Reports Server (NTRS)

Kramer, T.; Brogren, E.; Siegel, B.

1984-01-01

Cryogenic propellant tank insulations or liquid oxygen/liquid hydrogen low-thrust 2224N (500 lbf) propulsion systems (LTPS) were assessed. The insulation studied consisted of combinations of N2-purged foam and multilayer insulation (MLI) as well as He-purged MLI-only. Heat leak and payload performance predictions were made for three shuttle-launched LTPS designed for shuttle bay packaged payload densities of 56 kg cu/m (3.5 lbm/cu ft), 40 kg/cu m (2.5 lbm/cu ft) and 24 kg/cu m (1.5 lbm/cu ft). Foam/MLI insulations were found to increase LTPS payload delivery capability when compared with He-purged MLI-only. An additional benefit of foam/MLI was reduced operational complexity because orbiter cargo bay N2 purge gas could be used for MLI purging. Maximum payload mass benefit occurred when an enhanced convection, rather than natural convection, heat transfer was specified for the insulation purge enclosure. The enhanced convection environment allowed minimum insulation thickness to be used for the foam/MLI interface temperature selected to correspond to the moisture dew point in the N2 purge gas. Experimental verification of foam/MLI benefits was recommended. A conservative program cost estimate for testing a MLI-foam insulated tank was 2.1 million dollars. This cost could be reduced significantly without increasing program risk.

Entry Descent and Landing Workshop Proceedings. Volume 1; Inflatable Reentry Vehicle Experiment-3 (IRVE-3) Project Overview & Instrumentation

NASA Technical Reports Server (NTRS)

Dillman, Robert

2015-01-01

Entry mass at Mars is limited by the payload size that can be carried by a rigid capsule that can fit inside the launch vehicle fairing. Landing altitude at Mars is limited by ballistic coefficient (mass per area) of entry body. Inflatable technologies allow payload to use full diameter of launch fairing, and deploy larger aeroshell before atmospheric interface, landing more payload at a higher altitude. Also useful for return of large payloads from Low Earth Orbit (LEO).

Helicopter payload gains utilizing water injection for hot day power augmentation

NASA Technical Reports Server (NTRS)

Stroub, R. H.

1972-01-01

An analytical investigation was undertaken to assess the gains in helicopter mission payload through the use of water injection to produce power augmentation in an altitude-hot day environment. Substantial gains are shown for two representative helicopters, the UH-lH and CH-47B. The UH-lH payload increased 86.7 percent for a 50 n.mi. (92.6 km) radius mission involving two out-of-ground effect (OGE) hover take-offs of 2 minutes each at 5000 ft. (1525 m) 35 C ambient conditions. The CH-47B payload increased 49.5 percent for a 50 n.mi. (92.6 km) radius mission with sling loaded cargo as the outbound payload and a 3000 lb. (1360 kg) internal cargo on the return leg. The mission included two 4 min. OGE hovers at 6000 ft. (1830 m) 35 C. An improvement in take off performance and maximum performance climb also resulted as a consequence of the OGE hover capability and higher maximum power available.

Data base architecture for instrument characteristics critical to spacecraft conceptual design

NASA Technical Reports Server (NTRS)

Rowell, Lawrence F.; Allen, Cheryl L.

1990-01-01

Spacecraft designs are driven by the payloads and mission requirements that they support. Many of the payload characteristics, such as mass, power requirements, communication requirements, moving parts, and so forth directly affect the choices for the spacecraft structural configuration and its subsystem design and component selection. The conceptual design process, which translates mission requirements into early spacecraft concepts, must be tolerant of frequent changes in the payload complement and resource requirements. A computer data base was designed and implemented for the purposes of containing the payload characteristics pertinent for spacecraft conceptual design, tracking the evolution of these payloads over time, and enabling the integration of the payload data with engineering analysis programs for improving the efficiency in producing spacecraft designs. In-house tools were used for constructing the data base and for performing the actual integration with an existing program for optimizing payload mass locations on the spacecraft.

Payload specialists Millie Hughes-Fulford in Body Mass Measurement Device

NASA Image and Video Library

1985-02-01

S85-26553 (Feb 1985) --- STS-40/SLS-1 payload specialist Millie Hughes-Fulford sits strapped in the special device scientists have developed for determining mass on orbit. As the chair swings back and forth, a timer records how much the crewmember's mass retards the chair's movement. Dr. Hughes-Fulford will be joined by three mission specialists, the mission commander, the pilot and a second payload specialist for the scheduled 10-day Spacelab Life Sciences-1 (SLS-1) mission. The flight is totally dedicated to biological and medical experimentation.

14 CFR Appendix A to Part 440 - Information Requirements for Obtaining a Maximum Probable Loss Determination for Licensed or...

Code of Federal Regulations, 2010 CFR

2010-01-01

... payload, including type (e.g., telecommunications, remote sensing), propellants, and hazardous components... description of any payload, including type (e.g., telecommunications, remote sensing), propellants, and...

The unrealized potential for heavy balloon payloads

NASA Astrophysics Data System (ADS)

Winker, J. A.

1993-02-01

Knowing that properties of new polyethylene films are superior to previous types, one would believe that heavier payloads can now be safely carried. Some experimentation has already been done to verify that assumption. Future expectations are discussed. We believe that with present-day materials, and with only limited changes in design philosophies, maximum payload weights can be increased by 50 to 75% from presently accepted maxima.

Single-stage-to-orbit performance enhancement from take-off thrust augmentation

NASA Astrophysics Data System (ADS)

Galati, Terence; Elkins, Travis

1997-01-01

Thrust augmentation offers the Single Stage to Orbit (SSTO) space launch vehicle improved payload capability while reducing vehicle weight and cost. Optimization of vehicle configuration and flight profile are studied. Using a 612,000 kg Gross Lift Off Weight (GLOW) SSTO with three Castor® strap-on motors, payloads in excess of 18,000 kg to Low Earth Orbit (LEO) are achievable. Emphasis is placed on finding vehicle optimums in the 9,000 kg payload range to capture over 80% of commercial payloads. Strap-on boosters allow a small SSTO vehicle to fly with a mass fraction of only 0.88 and LOX/H2 engines operating at 445 sec vacuum specific impulse. Payload sensitivity due to variations of mass fraction, Isp and pitch rate are quantified.

Flight test of a spin parachute for use with a Super Arcas sounding rocket

NASA Technical Reports Server (NTRS)

Silbert, M. N.

1975-01-01

The development and flight testing of a specially configured 16.6 ft Disc Band Gap (DBG) Spin Parachute is discussed. The parachute is integrated with a modified Super Arcas launch vehicle. Total payload weight was 17.6 lbs including the Spin Parachute and a scientific payload, and lift-off weight was 100.3 lbs. The Super Arcas vehicle was despun from 18.4 cps. After payload separation at 244,170 ft the Spin Parachute and its payload attained a maximum spin rate of 2.4 cps. Total suspended weight of the Spin Parachute and its payload was 14.64 lbs.

Payload vehicle aerodynamic reentry analysis

NASA Astrophysics Data System (ADS)

Tong, Donald

An approach for analyzing the dynamic behavior of a cone-cylinder payload vehicle during reentry to insure proper deployment of the parachute system and recovery of the payload is presented. This analysis includes the study of an aerodynamic device that is useful in extending vehicle axial rotation through the maximum dynamic pressure region. Attention is given to vehicle configuration and reentry trajectory, the derivation of pitch static aerodynamics, the derivation of the pitch damping coefficient, pitching moment modeling, aerodynamic roll device modeling, and payload vehicle reentry dynamics. It is shown that the vehicle dynamics at parachute deployment are well within the design limit of the recovery system, thus ensuring successful payload recovery.

Performance and technical feasibility comparison of reusable launch systems: A synthesis of the ESA winged launcher studies

NASA Astrophysics Data System (ADS)

Berry, W.; Grallert, H.

1996-02-01

The paper presents a synthesis of the performance and technical feasibility assessment of 7 reusable launcher types, comprising 13 different vehicles, studied by European Industry for ESA in the ESA Winged Launcher Study in the period January 1988 to May 1994. The vehicles comprised single-stage-to-orbit (SSTO) and two-stage-to-orbit (TSTO) vehicles, propelled by either air-breathing/rocket propulsion or entirely by rocket propulsion. The results showed that an SSTO vehicle of the HOTOL-type, propelled by subsonic combustion air-breathing/rocket engines could barely deliver the specified payload mass and was aerodynamically unstable; that a TSTO vehicle of the Saenger type, employing subsonic combustion airbreathing propulsion in its first stage and rocket propulsion in its second stage, could readily deliver the specified payload mass and was found to be technically feasible and versatile; that an SSTO vehicle of the NASP type, propelled by supersonic combustion airbreathing/rocket propulsion was able to deliver a reduced payload mass, was very complex and required very advanced technologies; that an air-launched rocket propelled vehicle of the Interim HOTOL type, although technically feasible, could deliver only a reduced payload mass, being constrained by the lifting capability of the carrier airplane; that three different, entirely rocket-propelled vehicles could deliver the specified payload mass, were technically feasible but required relatively advanced technologies.

A neural network-based input shaping for swing suppression of an overhead crane under payload hoisting and mass variations

NASA Astrophysics Data System (ADS)

Ramli, Liyana; Mohamed, Z.; Jaafar, H. I.

2018-07-01

This paper proposes an improved input shaping for minimising payload swing of an overhead crane with payload hoisting and payload mass variations. A real time unity magnitude zero vibration (UMZV) shaper is designed by using an artificial neural network trained by particle swarm optimisation. The proposed technique could predict and directly update the shaper's parameters in real time to handle the effects of time-varying parameters during the crane operation with hoisting. To evaluate the performances of the proposed method, experiments are conducted on a laboratory overhead crane with a payload hoisting, different payload masses and two different crane motions. The superiority of the proposed method is confirmed by reductions of at least 38.9% and 91.3% in the overall and residual swing responses, respectively over a UMZV shaper designed using an average operating frequency and a robust shaper namely Zero Vibration Derivative-Derivative (ZVDD). The proposed method also demonstrates a significant residual swing suppression as compared to a ZVDD shaper designed based on varying frequency. In addition, the significant reductions are achieved with a less shaper duration resulting in a satisfactory speed of response. It is envisaged that the proposed method can be used for designing effective input shapers for payload swing suppression of a crane with time-varying parameters and for a crane that employ finite actuation states.

On locating steganographic payload using residuals

NASA Astrophysics Data System (ADS)

Quach, Tu-Thach

2011-02-01

Locating steganographic payload usingWeighted Stego-image (WS) residuals has been proven successful provided a large number of stego images are available. In this paper, we revisit this topic with two goals. First, we argue that it is a promising approach to locate payload by showing that in the ideal scenario where the cover images are available, the expected number of stego images needed to perfectly locate all load-carrying pixels is the logarithm of the payload size. Second, we generalize cover estimation to a maximum likelihood decoding problem and demonstrate that a second-order statistical cover model can be used to compute residuals to locate payload embedded by both LSB replacement and LSB matching steganography.

Large Space Optics: From Hubble to JWST and Beyond

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2008-01-01

If necessity truly is the mother of invention, then advances in lightweight space mirror technology have been driven by launch vehicle mass and volume constraints. In the late 1970 s, at the start of Hubble development, the state of the art in ground based telescopes was 3 to 4 meter monolithic primary mirrors with masses of 6000 to 10,000 kg - clearly too massive for the planned space shuttle 25,000 kg capability to LEO. Necessity led Hubble to a different solution. Launch vehicle mass constraints (and cost) resulted in the development of a 2.4 meter lightweight eggcrate mirror. At 810 kg (180 kg/m2), this mirror was approximately 7.4% of HST s total 11,110 kg mass. And, the total observatory structure at 4.3 m x 13.2 m fit snuggly inside the space shuttle 4.6 m x 18.3 m payload bay. In the early 1990 s, at the start of JWST development, the state of the art in ground based telescopes was 8 meter class monolithic primary mirrors (16,000 to 23,000 kg) and 10 meter segmented mirrors (14,400 kg). Unfortunately, launch vehicles were still constrained to 4.5 meter payloads and 25,000 kg to LEO or 6,600 kg to L2. Furthermore, science now demanded a space telescope with 6 to 8 meter aperture operating at L2. Mirror technology was identified as a critical capability necessary to enable the next generation of large aperture space telescopes. Specific telescope architectures were explored via three independent design concept studies conducted during the summer of 1996 (1). These studies identified two significant architectural constraints: segmentation and areal density. Because the launch vehicle fairing payload dynamic envelop diameter is approximately 4.5 meters, the only way to launch an 8 meter class mirror is to segment it, fold it and deploy it on orbit - resulting in actuation and control requirements. And, because of launch vehicle mass limits, the primary mirror allocation was only 1000 kg - resulting in a maximum areal density of 20 kg/m2. At the inception of JWST in 1996, such a capability did not exist. A highly successful technology development program was initiated resulting in matured and demonstrated mirror technology for JWST (2, 3). Today, the JWST 6.5 meter primary mirror has an areal density of 25 kg/m2 for a total mass of 625 kg or 9.6% of the total JWST observatory mass of 6,500 kg. Looking into the future, science requires increasing larger collecting apertures. Ground based telescopes are already moving towards 30+ meter mirrors. The only way to meet this challenge for space telescopes is via even lower areal density mirrors or on-orbit assembly or larger launch vehicles (4). The planned NASA Ares V with its 10 meter fairing and 55,000 kg payload to L2 eliminates this constraint (5).

Optimum data analysis procedures for Titan 4 and Space Shuttle payload acoustic measurements during lift-off

NASA Technical Reports Server (NTRS)

Piersol, Allan G.

1991-01-01

Analytical expressions have been derived to describe the mean square error in the estimation of the maximum rms value computed from a step-wise (or running) time average of a nonstationary random signal. These analytical expressions have been applied to the problem of selecting the optimum averaging times that will minimize the total mean square errors in estimates of the maximum sound pressure levels measured inside the Titan IV payload fairing (PLF) and the Space Shuttle payload bay (PLB) during lift-off. Based on evaluations of typical Titan IV and Space Shuttle launch data, it has been determined that the optimum averaging times for computing the maximum levels are (1) T (sub o) = 1.14 sec for the maximum overall level, and T(sub oi) = 4.88 f (sub i) (exp -0.2) sec for the maximum 1/3 octave band levels inside the Titan IV PLF, and (2) T (sub o) = 1.65 sec for the maximum overall level, and T (sub oi) = 7.10 f (sub i) (exp -0.2) sec for the maximum 1/3 octave band levels inside the Space Shuttle PLB, where f (sub i) is the 1/3 octave band center frequency. However, the results for both vehicles indicate that the total rms error in the maximum level estimates will be within 25 percent the minimum error for all averaging times within plus or minus 50 percent of the optimum averaging time, so a precise selection of the exact optimum averaging time is not critical. Based on these results, linear averaging times (T) are recommended for computing the maximum sound pressure level during lift-off.

Feasibility of Helicopter Support Seek Frost.

DTIC Science & Technology

1980-05-01

the allowable maximum weight can be used as the payload. The payload is a variable. Small helicopters with full fuel and auxillary tanks can fly...equipment, that the program to obtain icing approval on the S-76 will be finalized for management evaluation, and a decision can be made at that time to

Measurements of the Interaction of High-Voltage Biassed Conductors with the Ionosphere

DTIC Science & Technology

1989-01-01

tether from the mother payload. The daughter acts as the reference potential, and the electron beam is emitted from the mother . Sasaki et al. [1987...experiments, the maximum amount of charging of the mother payload was only on the order of 10 volts at 6 an altitude of 150 to 200 km. This low amount of...beam collisions with neutral atmospheric gases and that these electrons were the major current source to the mother payload, thus minimizing vehicle

European Space Agency /ESA/ pointing systems - Requirements and fulfillment

NASA Astrophysics Data System (ADS)

Nellessen, W.; Wolf, P.

1982-02-01

The Instrument Pointing System (IPS), a three-axes gimballed experimentation facility, is to be first used in the second Spacelab flight, in 1984. In response to crew motions and Space Shuttle thruster disturbances, payload mass and inertia, and IPS positioning in the cargo bay, the IPS will provide a stable pointing capability for payloads of up to 2000 kg whose accuracy is in the 1-10 arcsec range. The IPS is complemented in the up-to-200 kg payload mass range by a small Position and Hold Mount, a functional model of which is now available for performance demonstration tests.

Definition and design of an experiment to test raster scanning with rotating unbalanced-mass devices on gimbaled payloads

NASA Technical Reports Server (NTRS)

Lightsey, W. D.; Alhorn, D. C.; Polites, M. E.

1992-01-01

An experiment designed to test the feasibility of using rotating unbalanced-mass (RUM) devices for line and raster scanning gimbaled payloads, while expending very little power is described. The experiment is configured for ground-based testing, but the scan concept is applicable to ground-based, balloon-borne, and space-based payloads, as well as free-flying spacecraft. The servos used in scanning are defined; the electronic hardware is specified; and a computer simulation model of the system is described. Simulation results are presented that predict system performance and verify the servo designs.

High Pressure Earth Storable Rocket Technology Program: Basic Program

NASA Technical Reports Server (NTRS)

Chazen, M. L.; Sicher, D.; Huang, D.; Mueller, T.

1995-01-01

The HIPES Program was conducted for NASA-LeRC by TRW. The Basic Program consisted of system studies, design of testbed engine, fabrication and testing of engine. Studies of both pressure-fed and pump-fed systems were investigated for N2O4 and both MMH and N2H4 fuels with the result that N2H4 provides the maximum payload for all satellites over MMH. The higher pressure engine offers improved performance with smaller envelope and associated weight savings. Pump-fed systems offer maximum payload for large and medium weight satellites while pressure-fed systems offer maximum payload for small light weight satellites. The major benefits of HIPES are high performance within a confined length maximizing payload for lightsats which are length (volume) constrained. Three types of thrust chambers were evaluated -- Copper heatsink at 400, 500 and 600 psia chamber pressures for performance/thermal; water cooled to determine heat absorbed to predict rhenium engine operation; and rhenium to validate the concept. The HIPES engine demonstrated very high performance at 50 lbf thrust (epsilon = 150) and Pc = 500 psia with both fuels: Isp = 337 sec using N2O4-N2H4 and ISP = 327.5 sec using N2O4-MMH indicating combustion efficiencies greater than 98%. A powder metallurgy rhenium engine demonstrated operation with high performance at Pc = 500 psia which indicated the viability of the concept.

Advanced Propulsion for Geostationary Orbit Insertion and North-South Station Keeping

NASA Technical Reports Server (NTRS)

Oleson, Steven R.; Myers, Roger M.; Kluever, Craig A.; Riehl, John P.; Curran, Francis M.

1995-01-01

Solar electric propulsion (SEP) technology is currently being used for geostationary satellite station keeping to increase payload mass. Analyses show that advanced electric propulsion technologies can be used to obtain additional increases in payload mass by using these same technologies to perform part of the orbit transfer. In this work three electric propulsion technologies are examined at two power levels for an Atlas 2AS class spacecraft. The on-board chemical propulsion apogee engine fuel is reduced to allow the use of electric propulsion. A numerical optimizer is used to determine the chemical burns which will minimize the electric propulsion transfer time. Results show that for a 1550 kg Atlas 2AS class payload, increases in net mass (geostationary satellite mass less wet propulsion system mass) of 150 to 800 kg are possible using electric propulsion for station keeping, advanced chemical engines for part of the transfer, and electric propulsion for the remainder of the transfer. Trip times are between one and four months.

Systematic Propulsion Optimization Tools (SPOT)

NASA Technical Reports Server (NTRS)

Bower, Mark; Celestian, John

1992-01-01

This paper describes a computer program written by senior-level Mechanical Engineering students at the University of Alabama in Huntsville which is capable of optimizing user-defined delivery systems for carrying payloads into orbit. The custom propulsion system is designed by the user through the input of configuration, payload, and orbital parameters. The primary advantages of the software, called Systematic Propulsion Optimization Tools (SPOT), are a user-friendly interface and a modular FORTRAN 77 code designed for ease of modification. The optimization of variables in an orbital delivery system is of critical concern in the propulsion environment. The mass of the overall system must be minimized within the maximum stress, force, and pressure constraints. SPOT utilizes the Design Optimization Tools (DOT) program for the optimization techniques. The SPOT program is divided into a main program and five modules: aerodynamic losses, orbital parameters, liquid engines, solid engines, and nozzles. The program is designed to be upgraded easily and expanded to meet specific user needs. A user's manual and a programmer's manual are currently being developed to facilitate implementation and modification.

Ares V: Shifting the Payload Design Paradigm

NASA Technical Reports Server (NTRS)

Sumrall, Phil; Creech, Steve; Cockrell, Charles E.

2009-01-01

NASA is designing the Ares V heavy-lift cargo launch vehicle to send more crew and cargo to more places on the lunar surface than the 1960s-era Saturn V and to provide ongoing support for a permanent lunar outpost. This uncrewed cargo vehicle is designed to operate together with the Ares I crew vehicle (Figure 1). In addition to this role, however, its unmatched mass and volume capability represent a national asset for exploration, science, and commerce. The Ares V also enables or significantly enhances a large class of space missions not thought possible by scientists and engineers since the Saturn V program ended over 30 years ago. Compared to current systems, it will offer approximately five times the mass and volume to most orbits and locations. This should allow prospective mission planners to build robust payloads with margins that are three to five times the industry norm. The space inside the planned payload shroud has enough usable volume to launch the volumetric equivalent of approximately 10 Apollo Lunar Modules or approximately five equivalent Hubble Space Telescopes. This mass and volume capability to low-Earth orbit (LEO) enables a host of new scientific and observation platforms, such as telescopes, satellites, planetary and solar missions, as well as being able to provide the lift for future large in-space infrastructure missions, such as space based solar power and mining, Earth asteroid defense, propellant depots, etc. In addition, payload designers may also have the option of simplifying their designs or employing Ares V s payload as dumb mass to reduce technical and operational risk. The Ares V team is engaging the potential payload community now, two to three years before System Requirements Review (SRR), in order to better understand the additional requirements from the payload community that could be accommodated in the Ares V design in its conceptual phase. This paper will discuss the Ares V reference mission and capability, as well as its potential to perform other missions in the future.

Mass driver reaction engine characteristics and performance in earth orbital transfer missions

NASA Technical Reports Server (NTRS)

Snow, W. R.; Dunbar, R. S.

1982-01-01

Configurations of a typical mass driver reaction engine (MDRE) are presented and its use for delivery of payloads to geosynchronous orbit (GEO) from low earth orbit (LEO) is discussed. Basic rocket equations are developed for LEO to GEO round-trip missions using a single exhaust velocity. It is shown that exhaust velocities in the 5-10 km/sec range (specific impulse of 500-1000 sec) are well suited for mass drivers, minimizing the overall cost of missions. Payload delivery rate fractions show that there is little to be gained by stretching out LEO to GEO transfer times from 90 to 180 days. It therefore pays to use the shorter trip time, approximately doubling the amount of delivered payload during any fixed time of use of the MDRE.

High Altitude Launch for a Practical SSTO

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Denis, Vincent; Lyons, Valerie (Technical Monitor)

2003-01-01

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).

High Altitude Launch for a Practical SSTO

NASA Astrophysics Data System (ADS)

Landis, Geoffrey A.; Denis, Vincent

2003-01-01

Existing engineering materials allow the constuction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31% increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).

Atmosphere, Magnetosphere and Plasmas in Space (AMPS). Spacelab payload definition study. Volume 5: Technical summary

NASA Technical Reports Server (NTRS)

1976-01-01

Engineering and operational facets associated with the implementation of the first two AMPS flights are covered. The payload is described including all systems and subsystems and the mission planning and flight operations are described too. Payload integration, ground operations, and logistics are included along with key supporting analyses and mass properties.

Detecting Anomalies in Process Control Networks

NASA Astrophysics Data System (ADS)

Rrushi, Julian; Kang, Kyoung-Don

This paper presents the estimation-inspection algorithm, a statistical algorithm for anomaly detection in process control networks. The algorithm determines if the payload of a network packet that is about to be processed by a control system is normal or abnormal based on the effect that the packet will have on a variable stored in control system memory. The estimation part of the algorithm uses logistic regression integrated with maximum likelihood estimation in an inductive machine learning process to estimate a series of statistical parameters; these parameters are used in conjunction with logistic regression formulas to form a probability mass function for each variable stored in control system memory. The inspection part of the algorithm uses the probability mass functions to estimate the normalcy probability of a specific value that a network packet writes to a variable. Experimental results demonstrate that the algorithm is very effective at detecting anomalies in process control networks.

Space shuttle launch vehicle performance trajectory, exchange ratios, and dispersion analysis

NASA Technical Reports Server (NTRS)

Toelle, R. G.; Blackwell, D. L.; Lott, L. N.

1973-01-01

A baseline space shuttle performance trajectory for Mission 3A launched from WTR has been generated. Design constraints of maximum dynamic pressure, longitudinal acceleration, and delivered payload were satisfied. Payload exchange ratios are presented with explanation on use. Design envelopes of dynamic pressure, SRB staging point, aerodynamic heating and flight performance reserves are calculated and included.

Atmospheric, Magnetospheric, and Plasmas in Space (AMPS) spacelab payload definition study, appendixes

NASA Technical Reports Server (NTRS)

Keeley, J. T.

1976-01-01

An equipment list, instrument baseline data, engineering drawings, mass properties computer printouts, electrical energy management, and control and display functional analysis pertinent to the AMPS (Satellite Payload) are presented.

Using Paraffin PCM to Make Optical Communication Type of Payloads Thermally Self-Sufficient for Operation in Orion Crew Module

NASA Technical Reports Server (NTRS)

Choi, Michael K.

2016-01-01

An innovative concept of using paraffin phase change material with a melting point of 28 C to make Optical Communication type of payload thermally self-sufficient for operation in the Orion Crew Module is presented. It stores the waste heat of the payload and permits it to operate for about one hour by maintaining its temperature within the maximum operating limit. It overcomes the problem of relying on the availability of cold plate heat sink in the Orion Crew Module.

Project Icarus: Preliminary Thoughts on the Selection of Probes and Instruments for an Icarus-style Interstellar Mission

NASA Astrophysics Data System (ADS)

Crawford, Ian A.

2016-06-01

In this paper we outline the range of probes and scientific instruments that will be required in order for Icarus to fulfill its scientific mission of exploring a nearby star, its attendant planetary system, and the intervening interstellar medium. Based on this preliminary analysis, we estimate that the minimum total Icarus scientific payload mass (i.e. the mass of probes and instruments which must be decelerated to rest in the target system to enable a meaningful programme of scientific investigation) will be in the region of 100 tonnes. Of this, approximately 10 tonnes would be allocated for cruise-phase science instruments, and about 35 tonnes (i.e. the average of estimated lower and upper limits of 28 and 41 tonnes) would be contributed by the intra-system science payload itself (i.e. the dry mass of the stellar and planetary probes and their instruments). The remaining ~55 tonnes is allocated for the sub-probe intra-system propulsion requirements (crudely estimated from current Solar System missions; detailed modelling of sub-probe propulsion systems will be needed to refine this figure). The overall mass contributed by the science payload to the total that must be decelerated from the interstellar cruise velocity will be considerably more than 100 tonnes, however, as allowance must be made for the payload structural and infrastructural elements required to support, deploy, and communicate with the science probes and instruments. Based on the earlier Daedalus study, we estimate another factor of two to allow for these components. Pending the outcome of more detailed studies, it therefore appears that an overall science-related payload mass of ~200 tonnes will be required. This paper is a submission of the Project Icarus Study Group.

Cryo-braking using penetrators for enhanced capabilities for the potential landing of payloads on icy solar system objects

NASA Astrophysics Data System (ADS)

Winglee, R. M.; Robinson, T.; Danner, M.; Koch, J.

2018-03-01

The icy moons of Jupiter and Saturn are important astrobiology targets. Access to the surface of these worlds is made difficult by the high ΔV requirements which is typically in the hypervelocity range. Passive braking systems cannot be used due to the lack of an atmosphere, and active braking by rockets significantly adds to the missions costs. This paper demonstrates that a two-stage landing system can overcome these problems and provide significant improvements in the payload fraction that can be landed The first stage involves a hypervelocity impactor which is designed to penetrate to a depth of a few tens of meters. This interaction is the cryo-breaking component and is examined through laboratory experiments, empirical relations and modeling. The resultant ice-particle cloud creates a transient artificial atmosphere that can be used to enable passive braking of the second stage payload dd, with a substantially higher mass payload fraction than possible with a rocket landing system. It is shown that a hollow cylinder design for the impactor can more efficiently eject the material upwards in a solid cone of ice particles relative to solid impactors such as spheres or spikes. The ejected mass is shown to be of the order of 103 to 104 times the mass of the impactor. The modeling indicates that a 10 kg payload with a braking system of 3 m2 (i.e. an areal density of 0.3 kg/m2) is sufficient to allow the landing of the payload with the deceleration limited to less than 2000 g's. Modern electronics can withstand this deceleration and as such the system provides an important alternative to landing payloads on icy solar system objects.

Missile sizing for ascent-phase intercept

DOE Office of Scientific and Technical Information (OSTI.GOV)

Hull, D.G.; Salguero, D.E.

1994-11-01

A computer code has been developed to determine the size of a ground-launched, multistage missile which can intercept a theater ballistic missile before it leaves the atmosphere. Typical final conditions for the inteceptor are 450 km range, 60 km altitude, and 80 sec flight time. Given the payload mass (35 kg), which includes a kinetic kill vehicle, and achievable values for the stage mass fractions (0.85), the stage specific impulses (290 sec), and the vehicle density (60 lb/ft{sup 3}), the launch mass is minimized with respect to the stage payload mass ratios, the stage burn times, and the missile anglemore » of attack history subject to limits on the angle of attack (10 deg), the dynamic pressure (60,000 psf), and the maneuver load (200,000 psf deg). For a conical body, the minimum launch mass is approximately 1900 kg. The missile has three stages, and the payload coasts for 57 sec. A trade study has been performed by varying the flight time, the range, and the dynamic pressure Emits. With the results of a sizing study for a 70 lb payload and q{sub max} = 35,000 psf, a more detailed design has been carried out to determine heat shield mass, tabular aerodynamics, and altitude dependent thrust. The resulting missile has approximately 100 km less range than the sizing program predicted primarily because of the additional mass required for heat protection. On the other hand, launching the same missile from an aircraft increases its range by approximately 100 km. Sizing the interceptor for air launch with the same final conditions as the ground-launched missile reduces its launch mass to approximately 1000 kg.« less

Integrated payload and mission planning, phase 3. Volume 2: Logic/Methodology for preliminary grouping of spacelab and mixed cargo payloads

NASA Technical Reports Server (NTRS)

Rodgers, T. E.; Johnson, J. F.

1977-01-01

The logic and methodology for a preliminary grouping of Spacelab and mixed-cargo payloads is proposed in a form that can be readily coded into a computer program by NASA. The logic developed for this preliminary cargo grouping analysis is summarized. Principal input data include the NASA Payload Model, payload descriptive data, Orbiter and Spacelab capabilities, and NASA guidelines and constraints. The first step in the process is a launch interval selection in which the time interval for payload grouping is identified. Logic flow steps are then taken to group payloads and define flight configurations based on criteria that includes dedication, volume, area, orbital parameters, pointing, g-level, mass, center of gravity, energy, power, and crew time.

Heavy-Lift for a New Paradigm in Space Operations

NASA Technical Reports Server (NTRS)

Morris, Bruce; Burkey, Martin

2010-01-01

NASA is developing an unprecedented heavy-lift capability to enable human exploration beyond low Earth orbit (LEO). This capability could also significantly enhance numerous other missions of scientific, national security, and commercial importance. That capability is currently configured as the Ares V cargo launch vehicle. This capability will eclipse the capability the United States lost with the retirement of the Saturn V. It is capable of launching roughly 53 percent more payload mass to trans lunar injection (TLI) and 30 percent more payload mass to LEO than its Apollo Program predecessor. Ares V is a major element of NASA's Constellation Program, which also includes the Ares I crew launch vehicle (CLV), Orion crew exploration vehicle (CEV), and a lunar lander for crew and cargo. As currently configured, Ares V will be capable of launching 413,800 pounds (187.7 mT) to LEO, 138,500 pounds (63 mT) direct to the Moon or 156,700 pounds (71.1 mT) in its dual-launch architecture role with Ares I. Its 33-foot (10 m) shroud provides unprecedented payload volume. Assessment of astronomy and planetary science payload requirements since spring 2008 has indicated that a Saturn V-class heavy-lift vehicle has the potential to support a range of payloads and missions. This vehicle configuration enables some missions previously considered difficult or impossible and enhances many others. Collaborative design/architecture inputs, exchanges, and analyses have already begun between scientists and payload developers. This early dialogue between NASA engineers and payload designers allows both communities to shape their designs and operational concepts to be mutually supportive to the extent possible with the least financial impact. This paper provides an overview of the capabilities of a heavy-lift vehicle to launch payloads with increased mass and/or volume and reduce technical and cost risk in both design and operations.

A Minimal Radio and Plasma Wave Investigation For a Mercury Orbiter Mission

NASA Technical Reports Server (NTRS)

Kurth, W. S.

2001-01-01

The primary thrust of the effort at The University of Iowa for the definition of an orbiter mission to Mercury is a minimum viable radio and plasma wave investigation. While it is simple to add sensors and capability to any payload, the challenge is to do reasonable science within limited resources; and viable missions to Mercury are especially limited in payload mass. For a wave investigation, this is a serious concern, as the sensor mass often makes up a significant fraction of the instrumentation mass.

30-kW SEP Spacecraft as Secondary Payloads for Low-Cost Deep Space Science Missions

NASA Technical Reports Server (NTRS)

Brophy, John R.; Larson, Tim

2013-01-01

The Solar Array System contracts awarded by NASA's Space Technology Mission Directorate are developing solar arrays in the 30 kW to 50 kW power range (beginning of life at 1 AU) that have significantly higher specific powers (W/kg) and much smaller stowed volumes than conventional rigid-panel arrays. The successful development of these solar array technologies has the potential to enable new types of solar electric propulsion (SEP) vehicles and missions. This paper describes a 30-kW electric propulsion vehicle built into an EELV Secondary Payload Adapter (ESPA) ring. The system uses an ESPA ring as the primary structure and packages two 15-kW Megaflex solar array wings, two 14-kW Hall thrusters, a hydrazine Reaction Control Subsystem (RCS), 220 kg of xenon, 26 kg of hydrazine, and an avionics module that contains all of the rest of the spacecraft bus functions and the instrument suite. Direct-drive is used to maximize the propulsion subsystem efficiency and minimize the resulting waste heat and required radiator area. This is critical for packaging a high-power spacecraft into a very small volume. The fully-margined system dry mass would be approximately 1120 kg. This is not a small dry mass for a Discovery-class spacecraft, for example, the Dawn spacecraft dry mass was only about 750 kg. But the Dawn electric propulsion subsystem could process a maximum input power of 2.5 kW, and this spacecraft would process 28 kW, an increase of more than a factor of ten. With direct-drive the specific impulse would be limited to about 2,000 s assuming a nominal solar array output voltage of 300 V. The resulting spacecraft would have a beginning of life acceleration that is more than an order of magnitude greater than the Dawn spacecraft. Since the spacecraft would be built into an ESPA ring it could be launched as a secondary payload to a geosynchronous transfer orbit significantly reducing the launch costs for a planetary spacecraft. The SEP system would perform the escape from Earth and then the heliocentric transfer to the science target.

Launching rockets and small satellites from the lunar surface

NASA Technical Reports Server (NTRS)

Anderson, K. A.; Dougherty, W. M.; Pankow, D. H.

1985-01-01

Scientific payloads and their propulsion systems optimized for launch from the lunar surface differ considerably from their counterparts for use on earth. For spin-stabilized payloads, the preferred shape is a large diameter-to-length ratio to provide stability during the thrust phase. The rocket motor required for a 50-kg payload to reach an altitude of one lunar radius would have a mass of about 41 kg. To place spin-stabilized vehicles into low altitude circular orbits, they are first launched into an elliptical orbit with altitude about 840 km at aposelene. When the spacecraft crosses the desired circular orbit, small retro-rockets are fired to attain the appropriate direction and speed. Values of the launch angle, velocity increments, and other parameters for circular orbits of several altitudes are tabulated. To boost a 50-kg payload into a 100-km altitude circular orbit requires a total rocket motor mass of about 90 kg.

Launching rockets and small satellites from the lunar surface

NASA Astrophysics Data System (ADS)

Anderson, K. A.; Dougherty, W. M.; Pankow, D. H.

Scientific payloads and their propulsion systems optimized for launch from the lunar surface differ considerably from their counterparts for use on earth. For spin-stabilized payloads, the preferred shape is a large diameter-to-length ratio to provide stability during the thrust phase. The rocket motor required for a 50-kg payload to reach an altitude of one lunar radius would have a mass of about 41 kg. To place spin-stabilized vehicles into low altitude circular orbits, they are first launched into an elliptical orbit with altitude about 840 km at aposelene. When the spacecraft crosses the desired circular orbit, small retro-rockets are fired to attain the appropriate direction and speed. Values of the launch angle, velocity increments, and other parameters for circular orbits of several altitudes are tabulated. To boost a 50-kg payload into a 100-km altitude circular orbit requires a total rocket motor mass of about 90 kg.

Sounding rocket thermal analysis techniques applied to GAS payloads. [Get Away Special payloads (STS)

NASA Technical Reports Server (NTRS)

Wing, L. D.

1979-01-01

Simplified analytical techniques of sounding rocket programs are suggested as a means of bringing the cost of thermal analysis of the Get Away Special (GAS) payloads within acceptable bounds. Particular attention is given to two methods adapted from sounding rocket technology - a method in which the container and payload are assumed to be divided in half vertically by a thermal plane of symmetry, and a method which considers the container and its payload to be an analogous one-dimensional unit having the real or correct container top surface area for radiative heat transfer and a fictitious mass and geometry which model the average thermal effects.

Payload accommodation and development planning tools - A Desktop Resource Leveling Model (DRLM)

NASA Technical Reports Server (NTRS)

Hilchey, John D.; Ledbetter, Bobby; Williams, Richard C.

1989-01-01

The Desktop Resource Leveling Model (DRLM) has been developed as a tool to rapidly structure and manipulate accommodation, schedule, and funding profiles for any kind of experiments, payloads, facilities, and flight systems or other project hardware. The model creates detailed databases describing 'end item' parameters, such as mass, volume, power requirements or costs and schedules for payload, subsystem, or flight system elements. It automatically spreads costs by calendar quarters and sums costs or accommodation parameters by total project, payload, facility, payload launch, or program phase. Final results can be saved or printed out, automatically documenting all assumptions, inputs, and defaults.

CETF Space Station payload pointing system design and analysis feasibility study. [Critical Evaluation Task Force

NASA Technical Reports Server (NTRS)

Smagala, Tom; Mcglew, Dave

1988-01-01

The expected pointing performance of an attached payload coupled to the Critical Evaluation Task Force Space Station via a payload pointing system (PPS) is determined. The PPS is a 3-axis gimbal which provides the capability for maintaining inertial pointing of a payload in the presence of disturbances associated with the Space Station environment. A system where the axes of rotation were offset from the payload center of mass (CM) by 10 in. in the Z axis was studied as well as a system having the payload CM offset by only 1 inch. There is a significant improvement in pointing performance when going from the 10 in. to the 1 in. gimbal offset.

Mars Mobile Lander Systems for 2005 and 2007 Launch Opportunities

NASA Technical Reports Server (NTRS)

Sabahi, D.; Graf, J. E.

2000-01-01

A series of Mars missions are proposed for the August 2005 launch opportunity on a medium class Evolved Expendable Launch Vehicle (EELV) with a injected mass capability of 2600 to 2750 kg. Known as the Ranger class, the primary objective of these Mars mission concepts are: (1) Deliver a mobile platform to Mars surface with large payload capability of 150 to 450 kg (depending on launch opportunity of 2005 or 2007); (2) Develop a robust, safe, and reliable workhorse entry, descent, and landing (EDL) capability for landed mass exceeding 750 kg; (3) Provide feed forward capability for the 2007 opportunity and beyond; and (4) Provide an option for a long life telecom relay orbiter. A number of future Mars mission concepts desire landers with large payload capability. Among these concepts are Mars sample return (MSR) which requires 300 to 450 kg landed payload capability to accommodate sampling, sample transfer equipment and a Mars ascent vehicle (MAV). In addition to MSR, large in situ payloads of 150 kg provide a significant step up from the Mars Pathfinder (MPF) and Mars Polar Lander (MPL) class payloads of 20 to 30 kg. This capability enables numerous and physically large science instruments as well as human exploration development payloads. The payload may consist of drills, scoops, rock corers, imagers, spectrometers, and in situ propellant production experiment, and dust and environmental monitoring.

High Altitude Launch for a Practical SSTO

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Denis, Vincent

2003-01-01

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. High-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2 percent of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20 percent increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2 percent of gross lift-off weight, this corresponds to 31 percent increase in payload (for 5-kilometer launch altitude) to 122 percent additional payload (for 25-kilometer launch altitude).

High Altitude Launch for a Practical SSTO

NASA Technical Reports Server (NTRS)

Landis, Geoffrey A.; Denis, Vincent

2003-01-01

Existing engineering materials allow the construction of towers to heights of many kilometers. Orbital launch from a high altitude has significant advantages over sea-level launch due to the reduced atmospheric pressure, resulting in lower atmospheric drag on the vehicle and allowing higher rocket engine performance. high-altitude launch sites are particularly advantageous for single-stage to orbit (SSTO) vehicles, where the payload is typically 2% of the initial launch mass. An earlier paper enumerated some of the advantages of high altitude launch of SSTO vehicles. In this paper, we calculate launch trajectories for a candidate SSTO vehicle, and calculate the advantage of launch at launch altitudes 5 to 25 kilometer altitudes above sea level. The performance increase can be directly translated in to increased payload capability to orbit, ranging from 5 to 20% increase in the mass to orbit. For a candidate vehicle with an initial payload fraction of 2% of gross lift-off weight, this corresponds to 31 % increase in payload (for 5-km launch altitude) to 122% additional payload (for 25-km launch altitude).

Approach to Spacelab Payload mission management

NASA Technical Reports Server (NTRS)

Craft, H. G.; Lester, R. C.

1978-01-01

The nucleus of the approach to Spacelab Payload mission management is the establishment of a single point of authority for the entire payload on a given mission. This single point mission manager will serve as a 'broker' between the individual experiments and the STS, negotiating agreements by two-part interaction. The payload mission manager, along with a small support team, will represent the users in negotiating use of STS accommodations. He will provide the support needed by each individual experimenter to meet the scientific, technological, and applications objectives of the mission with minimum cost and maximum efficiency. The investigator will assume complete responsibility for his experiment hardware definition and development and will take an active role in the integration and operation of his experiment.

SLS Overview and Progress

NASA Technical Reports Server (NTRS)

Honeycutt, John

2017-01-01

Space Launch System will be able to offer payload accommodations with five times more volume than any contemporary launch vehicle Payload fairings of up to 10-meter diameter are being studied Space Launch System will offer an initial capability of greater than 70 metric tons to low Earth orbit; current U.S. launch vehicle maximum is 28 t Evolved version of SLS will offer Mars-enabling capability of greater than 130 metric tons to LEO SLS offers reduced transit times to the outer solar system by half or greater Higher characteristic energy (C3) also enables larger payloads to destination

LAVA subsystem integration and testing for the Resolve payload of the Resource Prospector mission: mass spectrometers and gas chromatography

NASA Astrophysics Data System (ADS)

Stewart, Elaine M.; Coan, Mary R.; Captain, Janine; Santiago-Bond, Josephine

2016-09-01

In-Situ Resource Utilization (ISRU) is a key NASA initiative to exploit resources at the site of planetary exploration for mission-critical consumables, propellants, and other supplies. The Resource Prospector mission, part of ISRU, is scheduled to launch in 2020 and will include a rover and lander hosting the Regolith and Environment Science and Oxygen and Lunar Volatile Extraction (RESOLVE) payload for extracting and analyzing lunar resources, particularly low molecular weight volatiles for fuel, air, and water. RESOLVE contains the Lunar Advanced Volatile Analysis (LAVA) subsystem with a Gas Chromatograph-Mass Spectrometer (GC-MS). RESOLVE subsystems, including the RP15 rover and LAVA, are in NASA's Engineering Test Unit (ETU) phase to assure that all vital components of the payload are space-flight rated and will perform as expected during the mission. Integration and testing of LAVA mass spectrometry verified reproducibility and accuracy of the candidate MS for detecting nitrogen, oxygen, and carbon dioxide. The RP15 testing comprised volatile analysis of water-doped simulant regolith to enhance integration of the RESOLVE payload with the rover. Multiple tests show the efficacy of the GC to detect 2% and 5% water-doped samples.

Preliminary design of the thermal protection system for solar probe

NASA Technical Reports Server (NTRS)

Dirling, R. B., Jr.; Loomis, W. C.; Heightland, C. N.

1982-01-01

A preliminary design of the thermal protection system for the NASA Solar Probe spacecraft is presented. As presently conceived, the spacecraft will be launched by the Space Shuttle on a Jovian swing-by trajectory and at perihelion approach to three solar radii of the surface of the Earth's sun. The system design satisfies maximum envelope, structural integrity, equipotential, and mass loss/contamination requirements by employing lightweight carbon-carbon emissive shields. The primary shield is a thin shell, 15.5-deg half-angle cone which absorbs direct solar flux at up to 10-deg off-nadir spacecraft pointing angles. Secondary shields of sandwich construction and low thickness-direction thermal conductivity are used to reduce the primary shield infrared radiation to the spacecraft payload.

Performance of a recoverable tug for planetary missions including use of perigee propulsion and corrections for nodal regression

NASA Technical Reports Server (NTRS)

Borsody, J.

1976-01-01

Mathematical equations are derived by using the Maximum Principle to obtain the maximum payload capability of a reusable tug for planetary missions. The mathematical formulation includes correction for nodal precession of the space shuttle orbit. The tug performs this nodal correction in returning to this precessed orbit. The sample case analyzed represents an inner planet mission as defined by the declination (fixed) and right ascension of the outgoing asymptote and the mission energy. Payload capability is derived for a typical cryogenic tug and the sample case with and without perigee propulsion. Optimal trajectory profiles and some important orbital elements are also discussed.

Passive orientation apparatus

DOEpatents

Spletzer, Barry L.; Fischer, Gary J.; Martinez, Michael A.

2001-01-01

An apparatus that can return a payload to a known orientation after unknown motion, without requiring external power or complex mechanical systems. The apparatus comprises a faceted cage that causes the system to rest in a stable position and orientation after arbitrary motion. A gimbal is mounted with the faceted cage and holds the payload, allowing the payload to move relative to the stable faceted cage. The payload is thereby placed in a known orientation by the interaction of gravity with the geometry of the faceted cage, the mass of the system, and the motion of the payload and gimbal. No additional energy, control, or mechanical actuation is required. The apparatus is suitable for use in applications requiring positioning of a payload to a known orientation after arbitrary or uncontrolled motion, including remote sensing and mobile robot applications.

COSMO - SkyMed Mission Overview

DTIC Science & Technology

2000-10-01

antenna with range and cross-range steering capabilities; The SAR Payload is an X-band Radar which 0 development and qualification of low mass ...summarised as follows: * Swaths: 20 Kmn to 300 Km SIlfale sflos" Swaccs: regKion: t 350 (a) to support the Payload mass (on ground,"• Access region: -/+ 35...real-time product is requested); situ" product delivery. This raises the problem of the size of the data to be transmitted and the geo- Customisation

Development of a Solar Array Drive Assembly for CubeSat

NASA Technical Reports Server (NTRS)

Passaretti, Mike; Hayes, Ron

2010-01-01

Small satellites and in particular CubeSats, have increasingly become more viable as platforms for payloads typically requiring much larger bus structures. As advances in technology make payloads and instruments for space missions smaller, lighter and more power efficient, a niche market is emerging from the university community to perform rapidly developed, low-cost missions on very small spacecraft - micro, nano, and picosatellites. In just the last few years, imaging, biological and new technology demonstration missions have been either proposed or have flown using variations of the CubeSat structure as a basis. As these missions have become more complex, and the CubeSat standard has increased in both size (number of cubes) and mass, available power has become an issue. Body-mounted solar cells provide a minimal amount of power; deployable arrays improve on that baseline but are still limited. To truly achieve maximum power, deployed tracked arrays are necessary. To this end, Honeybee Robotics Spacecraft Mechanisms Corporation, along with MMA of Nederland Colorado, has developed a solar array drive assembly (SADA) and deployable solar arrays specifically for CubeSat missions. In this paper, we discuss the development of the SADA.

Analytical Assessment of the Reciprocating Feed System

NASA Technical Reports Server (NTRS)

Eddleman, David E.; Blackmon, James B.; Morton, Christopher D.

2006-01-01

A preliminary analysis tool has been created in Microsoft Excel to determine deliverable payload mass, total system mass, and performance of spacecraft systems using various types of propellant feed systems. These mass estimates are conducted by inserting into the user interface the basic mission parameters (e.g., thrust, burn time, specific impulse, mixture ratio, etc.), system architecture (e.g., propulsion system type and characteristics, propellants, pressurization system type, etc.), and design properties (e.g., material properties, safety factors, etc.). Different propellant feed and pressurization systems are available for comparison in the program. This gives the user the ability to compare conventional pressure fed, reciprocating feed system (RFS), autogenous pressurization thrust augmentation (APTA RFS), and turbopump systems with the deliverable payload, inert mass, and total system mass being the primary comparison metrics. Analyses of several types of missions and spacecraft were conducted and it was found that the RFS offers a performance improvement, especially in terms of delivered payload, over conventional pressure fed systems. Furthermore, it is competitive with a turbopump system at low to moderate chamber pressures, up to approximately 1,500 psi. Various example cases estimating the system mass and deliverable payload of several types of spacecraft are presented that illustrate the potential system performance advantages of the RFS. In addition, a reliability assessment of the RFS was conducted, comparing it to simplified conventional pressure fed and turbopump systems, based on MIL-STD 756B; these results showed that the RFS offers higher reliability, and thus substantially longer periods between system refurbishment, than turbopump systems, and is competitive with conventional pressure fed systems. This is primarily the result of the intrinsic RFS fail-operational capability with three run tanks, since the system can operate with just two run tanks.

STS-97 P6 truss payload canister is lifted into payload changeout room

NASA Technical Reports Server (NTRS)

2000-01-01

On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, is lifted toward the payload changeout room (PCR). The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure that supports payload delivery at the pad and subsequent vertical installation in the orbiter payload bay. Attached to the canister are the red umbilical lines that maintain the controlled environment inside. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station'''s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a '''blanket''' that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST.

Project LOCOST: Laser or Chemical Hybrid Orbital Space Transport

NASA Technical Reports Server (NTRS)

Dixon, Alan; Kost, Alicia; Lampshire, Gregory; Larsen, Rob; Monahan, Bob; Wright, Geoff

1990-01-01

A potential mission in the late 1990s is the servicing of spacecraft assets located in GEO. The Geosynchronous Operations Support Center (GeoShack) will be supported by a space transfer vehicle based at the Space Station (SS). The vehicle will transport cargo between the SS and the GeoShack. A proposed unmanned, laser or chemical hybrid orbital space transfer vehicle (LOCOST) can be used to efficiently transfer cargo between the two orbits. A preliminary design shows that an unmanned, laser/chemical hybrid vehicle results in the fuel savings needed while still providing fast trip times. The LOCOST vehicle receives a 12 MW laser beam from one Earth orbiting, solar pumped, iodide Laser Power Station (LPS). Two Energy Relay Units (ERU) provide laser beam support during periods of line-of-sight blockage by the Earth. The baseline mission specifies a 13 day round trip transfer time. The ship's configuration consist of an optical train, one hydrogen laser engine, two chemical engines, a 18 m by 29 m box truss, a mission-flexible payload module, and propellant tanks. Overall vehicle dry mass is 8,000 kg. Outbound cargo mass is 20,000 kg, and inbound cargo mass is 6,000 kg. The baseline mission needs 93,000 kg of propellants to complete the scenario. Fully fueled, outbound mission mass is 121,000 kg. A regeneratively cooled, single plasma, laser engine design producing a maximum of 768 N of thrust is utilized along with two traditional chemical engines. The payload module is designed to hold 40,000 kg of cargo, though the baseline mission specifies less. A proposed design of a laser/chemical hybrid vehicle provides a trip time and propellant efficient means to transport cargo from the SS to a GeoShack. Its unique, hybrid propulsion system provides safety through redundancy, allows baseline missions to be efficiently executed, while still allowing for the possibility of larger cargo transfers.

NASA's Space Launch System: Progress Toward the Proving Ground

NASA Technical Reports Server (NTRS)

Jackman, Angie

2017-01-01

Space Launch System will be able to offer payload accommodations with five times more volume than any contemporary launch vehicle. center dot Payload fairings of up to 10-meter diameter are planned. Space Launch System will offer an initial capability of greater than 70 metric tons to low Earth orbit; current U.S. launch vehicle maximum is 28 t. center dot Evolved version of SLS will offer greatest-ever capability of greater than 130 t to LEO. SLS offers reduced transit times to the outer solar system by half or greater. center dot Higher characteristic energy (C3) also enables larger payloads to destination.

Optimization of transfer trajectories to the Apophis asteroid for spacecraft with high and low thrust

NASA Astrophysics Data System (ADS)

Ivashkin, V. V.; Krylov, I. V.

2014-03-01

The problem of optimization of a spacecraft transfer to the Apophis asteroid is investigated. The scheme of transfer under analysis includes a geocentric stage of boosting the spacecraft with high thrust, a heliocentric stage of control by a low thrust engine, and a stage of deceleration with injection to an orbit of the asteroid's satellite. In doing this, the problem of optimal control is solved for cases of ideal and piecewise-constant low thrust, and the optimal magnitude and direction of spacecraft's hyperbolic velocity "at infinity" during departure from the Earth are determined. The spacecraft trajectories are found based on a specially developed comprehensive method of optimization. This method combines the method of dynamic programming at the first stage of analysis and the Pontryagin maximum principle at the concluding stage, together with the parameter continuation method. The estimates are obtained for the spacecraft's final mass and for the payload mass that can be delivered to the asteroid using the Soyuz-Fregat carrier launcher.

The Pluto fast flyby mission: Completing the reconnaissance of the solar system

NASA Technical Reports Server (NTRS)

Henry, Paul K.

1993-01-01

The concept of a fast flyby mission to Pluto has been advanced as a means to complete the reconnaissance of the known solar system. In order to acquire data on the Pluto system at the earliest possible time, and within the professional lifetime of investigators now active in the field, concepts are being developed for relatively small spacecraft in the mass range of 70 Kg to 350 Kg with flight times to Pluto of 7 to 13 years. Necessarily, the science complement on such a mission will be very mass and power limited. The challenge will be to define a spacecraft and an instrument package that will maximize the scientific return within these limitations. Cost, of course, will be a major consideration, and funds for new technology development specific to this mission will not be extensive. Consequently, innovative ways to incorporate elegant simplicity into the designs must be found. In order to facilitate exploration of the Pluto-Charon system, fully integrated science payloads must be developed. Two proposed mission designs involving limited mass and power science payloads have been presented to the Outer Planets Science Working Group (OPSWG). These payload mass allocations range from 5 to 30 kilograms with power allocations as low as 5 watts. The drivers behind these low mass and power allocations are that they enable developing missions to fit within the moderate mission cost profile and allow fast flight times to Pluto (7 to 13 years). The OPSWG has prioritized science goals for this class of reconnaissance mission. Three specific science objectives were identified as the highest priority required for the first Pluto mission. These goals were: (1) study of the neutral atmosphere, (2) geology and morphology, and (3) surface compositional mapping. In order to achieve these science goals within the constraints of low mass, power and cost, it may be necessary to combine the functions of 3 conventional instruments (CCD camera, Ultra-Violet Spectrometer, and Infrared Spectrometer) into one fully integrated payload. Where possible, this payload would share optics, mechanisms, electronics and packaging.

Design of Sounding Rocket Payloads.

DTIC Science & Technology

1981-07-01

AD-AlB 271 NORTHEASTERN UNIV BOSTON MASS ELECTRONICS RESEARCH LAB F/6 19/7 DESIGN OF SOUNDING ROCKET PAYLOADS. (U) JUL Al R L MORIN, L .J O’CONNOR...Morin Lawrence J. O’Connor NORTHEASTERN UNIVERSITY Electronics Research Laboratory D T I Boston, Massachusetts 02115 ELECTE S DEC 9 19813 FINAL REPORT... Research Range on 21 February 1978. The payload was re-assembled, checked and mated to the launch vehicle on 27 February. Launch -8- criteria were

Alternative Way of Shifting Mass to Move a Spherical Robot

NASA Technical Reports Server (NTRS)

Lux, James

2005-01-01

The method proposed calls for suspending a payload by use of four or more cables that would be anchored to the inner surface of the sphere. In this method, the anchor points would not be diametrically opposite points defining Cartesian axes. The payload, which includes the functional analog of the aforementioned control box, would contain winches that would shorten or lengthen the cables in a coordinated manner to shift the position of the payload within the shell.

Benefits of slush hydrogen for space missions

NASA Technical Reports Server (NTRS)

Friedlander, Alan; Zubrin, Robert; Hardy, Terry L.

1991-01-01

A study was performed to quantify the benefits of using slush hydrogen instead of normal boiling point liquid hydrogen as a fuel for several space missions. Vehicles considered in the study included the Space Shuttle/Shuttle-C, LEO to GEO transfer vehicles, Lunar and Mars transfer vehicles, and cryogenic depots in low Earth orbit. The advantages of using slush hydrogen were expressed in terms of initial mass differences at a constant payload, payload differences at a constant tank volume, and increases in fuel storage time for cryogenic depots. Both chemical oxygen/hydrogen and hydrogen nuclear thermal rocket propulsion were considered in the study. The results indicated that slush hydrogen offers the potential for significant decreases in initial mass and increases in payload for most missions studied. These advantages increase as the mission difficulty, or energy, increases.

Conformational Assessment of Adnectin and Adnectin-Drug Conjugate by Hydrogen/Deuterium Exchange Mass Spectrometry

NASA Astrophysics Data System (ADS)

Huang, Richard Y.-C.; O'Neil, Steven R.; Lipovšek, Daša; Chen, Guodong

2018-05-01

Higher-order structure (HOS) characterization of therapeutic protein-drug conjugates for comprehensive assessment of conjugation-induced protein conformational changes is an important consideration in the biopharmaceutical industry to ensure proper behavior of protein therapeutics. In this study, conformational dynamics of a small therapeutic protein, adnectin 1, together with its drug conjugate were characterized by hydrogen/deuterium exchange mass spectrometry (HDX-MS) with different spatial resolutions. Top-down HDX allows detailed assessment of the residue-level deuterium content in the payload conjugation region. HDX-MS dataset revealed the ability of peptide-based payload/linker to retain deuterium in HDX experiments. Combined results from intact, top-down, and bottom-up HDX indicated no significant conformational changes of adnectin 1 upon payload conjugation. [Figure not available: see fulltext.

Design, fabrication & performance analysis of an unmanned aerial vehicle

NASA Astrophysics Data System (ADS)

Khan, M. I.; Salam, M. A.; Afsar, M. R.; Huda, M. N.; Mahmud, T.

2016-07-01

An Unmanned Aerial Vehicle was designed, analyzed and fabricated to meet design requirements and perform the entire mission for an international aircraft design competition. The goal was to have a balanced design possessing, good demonstrated flight handling qualities, practical and affordable manufacturing requirements while providing a high vehicle performance. The UAV had to complete total three missions named ferry flight (1st mission), maximum load mission (2nd mission) and emergency medical mission (3rd mission). The requirement of ferry flight mission was to fly as many as laps as possible within 4 minutes. The maximum load mission consists of flying 3 laps while carrying two wooden blocks which simulate cargo. The requirement of emergency medical mission was complete 3 laps as soon as possible while carrying two attendances and two patients. A careful analysis revealed lowest rated aircraft cost (RAC) as the primary design objective. So, the challenge was to build an aircraft with minimum RAC that can fly fast, fly with maximum payload, and fly fast with all the possible configurations. The aircraft design was reached by first generating numerous design concepts capable of completing the mission requirements. In conceptual design phase, Figure of Merit (FOM) analysis was carried out to select initial aircraft configuration, propulsion, empennage and landing gear. After completion of the conceptual design, preliminary design was carried out. The preliminary design iterations had a low wing loading, high lift coefficient, and a high thrust to weight ratio. To make the aircraft capable of Rough Field Taxi; springs were added in the landing gears for absorbing shock. An airfoil shaped fuselage was designed to allowed sufficient space for payload and generate less drag to make the aircraft fly fast. The final design was a high wing monoplane with conventional tail, single tractor propulsion system and a tail dragger landing gear. Payload was stored in undercarriage box for maximum load mission and emergency medical mission. The aircraft structure, weights 5.6 lb., constructed by balsa wood, depron and covering film was the only feasible match for the given requirements set by the competition organizers. The defined final aircraft was capable of: Completing 3 laps within 4 minutes at the first mission; flying 3 laps with 4 internal payloads at the second mission; flying 3 laps with all possible payload configurations at the third mission.

Increasing the Endurance and Payload Capacity of Unmanned Vehicles with Thin-Film Photovoltaics

DTIC Science & Technology

2014-06-01

25  1.  Maximum Power Point Tracker .......................................................25  2.  DC-DC Power Conversion...An example of the amorphous silicon cell (from [28]). ...................................20  Figure 15.  The structure of a copper indium gallium ...for the final solar array with the maximum power point indicated

Compact tunable and reconfigurable microwave photonic filter for satellite payloads

NASA Astrophysics Data System (ADS)

Santos, M. C.; Yoosefi, O.

2017-11-01

The trend towards the photonic processing of electrical signals at microwave frequencies for satellite payloads is increasing at a breathtaking pace, mainly spurred by prospects of wide electrical bandwidth operation, low mass and volume, reduced electrical noise levels, immunity to electromagnetic interferences and resistance to both temperature and radiation.

Space Launch System for Exploration and Science

NASA Astrophysics Data System (ADS)

Klaus, K.

2013-12-01

Introduction: The Space Launch System (SLS) is the most powerful rocket ever built and provides a critical heavy-lift launch capability enabling diverse deep space missions. The exploration class vehicle launches larger payloads farther in our solar system and faster than ever before. The vehicle's 5 m to 10 m fairing allows utilization of existing systems which reduces development risks, size limitations and cost. SLS lift capacity and superior performance shortens mission travel time. Enhanced capabilities enable a myriad of missions including human exploration, planetary science, astrophysics, heliophysics, planetary defense and commercial space exploration endeavors. Human Exploration: SLS is the first heavy-lift launch vehicle capable of transporting crews beyond low Earth orbit in over four decades. Its design maximizes use of common elements and heritage hardware to provide a low-risk, affordable system that meets Orion mission requirements. SLS provides a safe and sustainable deep space pathway to Mars in support of NASA's human spaceflight mission objectives. The SLS enables the launch of large gateway elements beyond the moon. Leveraging a low-energy transfer that reduces required propellant mass, components are then brought back to a desired cislunar destination. SLS provides a significant mass margin that can be used for additional consumables or a secondary payloads. SLS lowers risks for the Asteroid Retrieval Mission by reducing mission time and improving mass margin. SLS lift capacity allows for additional propellant enabling a shorter return or the delivery of a secondary payload, such as gateway component to cislunar space. SLS enables human return to the moon. The intermediate SLS capability allows both crew and cargo to fly to translunar orbit at the same time which will simplify mission design and reduce launch costs. Science Missions: A single SLS launch to Mars will enable sample collection at multiple, geographically dispersed locations and a low-risk, direct return of Martian material. For the Europa Clipper mission the SLS eliminates Venus and Earth flybys, providing a direct launch to the Jovian system, arriving four years earlier than missions utilizing existing launch vehicles. This architecture allows increased mass for radiation shielding, expansion of the science payload and provides a model for other outer planet missions. SLS provides a direct launch to the Uranus system, reducing travel time by two years when compared to existing launch capabilities. SLS can launch the Advanced Technology Large-Aperture Space Telescope (ATLAST 16 m) to SEL2, providing researchers 10 times the resolution of the James Webb Space Telescope and up to 300 times the sensitivity of the Hubble Space Telescope. SLS is the only vehicle capable of deploying telescopes of this mass and size in a single launch. It simplifies mission design and reduces risks by eliminating the need for multiple launches and in-space assembly. SLS greatly shortens interstellar travel time, delivering the Interstellar Explorer to 200 AU in about 15 years with a maximum speed of 63 km/sec--13.3 AU per year (Neptune orbits the sun at an approximate distance of 30 AU ).

Preliminary Findings of the Photovoltaic Cell Calibration Experiment on Pathfinder Flight 95-3

NASA Technical Reports Server (NTRS)

Vargas-Aburto, Carlos

1997-01-01

The objective of the photovoltaic (PV) cell calibration experiment for Pathfinder was to develop an experiment compatible with an ultralight UAV to predict the performance of PV cells at AM0, the solar spectrum in space, using the Langley plot technique. The Langley plot is a valuable technique for this purpose and requires accurate measurements of air mass (pressure), cell temperature, solar irradiance, and current-voltage(IV) characteristics with the cells directed normal to the direct ray of the sun. Pathfinder's mission objective (95-3) of 65,000 ft. maximum altitude, is ideal for performing the Langley plot measurements. Miniaturization of electronic data acquisition equipment enabled the design and construction of an accurate and light weight measurement system that meets Pathfinder's low payload weight requirements.

Sensitivity Analysis of Hybrid Propulsion Transportation System for Human Mars Expeditions

NASA Technical Reports Server (NTRS)

Chai, Patrick R.; Joyce, Ryan T.; Kessler, Paul D.; Merrill, Raymond G.; Qu, Min

2017-01-01

The National Aeronautics and Space Administration continues to develop and refine various transportation options to successfully field a human Mars campaign. One of these transportation options is the Hybrid Transportation System which utilizes both solar electric propulsion and chemical propulsion. The Hybrid propulsion system utilizes chemical propulsion to perform high thrust maneuvers, where the delta-V is most optimal when ap- plied to save time and to leverage the Oberth effect. It then utilizes solar electric propulsion to augment the chemical burns throughout the interplanetary trajectory. This eliminates the need for the development of two separate vehicles for crew and cargo missions. Previous studies considered single point designs of the architecture, with fixed payload mass and propulsion system performance parameters. As the architecture matures, it is inevitable that the payload mass and the performance of the propulsion system will change. It is desirable to understand how these changes will impact the in-space transportation system's mass and power requirements. This study presents an in-depth sensitivity analysis of the Hybrid crew transportation system to payload mass growth and solar electric propulsion performance. This analysis is used to identify the breakpoints of the current architecture and to inform future architecture and campaign design decisions.

Spacecraft configuration study for second generation mobile satellite system

NASA Technical Reports Server (NTRS)

Louie, M.; Vonstentzsch, W.; Zanella, F.; Hayes, R.; Mcgovern, F.; Tyner, R.

1985-01-01

A high power, high performance communicatons satellite bus being developed is designed to satisfy a broad range of multimission payload requirements in a cost effective manner and is compatible with both STS and expendable launchers. Results are presented of tradeoff studies conducted to optimize the second generation mobile satellite system for its mass, power, and physical size. Investigations of the 20-meter antenna configuration, transponder linearization techniques, needed spacecraft modifications, and spacecraft power, dissipation, mass, and physical size indicate that the advanced spacecraft bus is capable of supporting the required payload for the satellite.

Mercury orbiter transport study

NASA Technical Reports Server (NTRS)

Friedlander, A. L.; Feingold, H.

1977-01-01

A data base and comparative performance analyses of alternative flight mode options for delivering a range of payload masses to Mercury orbit are provided. Launch opportunities over the period 1980-2000 are considered. Extensive data trades are developed for the ballistic flight mode option utilizing one or more swingbys of Venus. Advanced transport options studied include solar electric propulsion and solar sailing. Results show the significant performance tradeoffs among such key parameters as trip time, payload mass, propulsion system mass, orbit size, launch year sensitivity and relative cost-effectiveness. Handbook-type presentation formats, particularly in the case of ballistic mode data, provide planetary program planners with an easily used source of reference information essential in the preliminary steps of mission selection and planning.

Radar cross-section measurements and simulation of a tethered satellite. The small expendable deployer system end-mass payload

NASA Technical Reports Server (NTRS)

Cravey, Robin L.; Fralick, Dion T.; Vedeler, Erik

1995-01-01

The first Small Expendable Deployer System (SEDS-1), a tethered satellite system, was developed by NASA and launched March 29, 1993 as a secondary payload on a United State Air Force (USAF) Delta-2 launch vehicle. The SEDS-1 successfully deployed an instrumented end-mass payload (EMP) on a 20-km nonconducting tether from the second stage of the Delta 2. This paper describes the effort of NASA Langley Research Center's Antenna and Microwave Research Branch to provide assistance to the SEDS Investigators Working Group (IWG) in determining EMP dynamics by analyzing the mission radar skin track data. The radar cross section measurements taken and simulations done for this study are described and comparisons of the measured data with the simulated data for the EMP at 6 GHz are presented.

Defense Science Board Summer Study on Autonomy

DTIC Science & Technology

2016-06-01

hours, at a maximum velocity of 40 mph, with a maximum payload of 9 kg (20 lbs); a maximum range of 160 km (100 miles); and can operate in wind /gust...existing mine disposal platform, such as Seafox, with contact reacquisition and neutralization capability. Seafox is a wire -guided mine neutralizer...functions, will retain operator control of neutralization and will remove the need for personnel to enter the minefield to execute fly- by- wire

Results of a laboratory experiment that tests rotating unbalanced-mass devices for scanning gimbaled payloads and free-flying spacecraft

NASA Technical Reports Server (NTRS)

Alhorn, D. C.; Polites, M. E.

1994-01-01

Rotating unbalanced-mass (RUM) devices are a new way to scan space-based, balloon-borne, and ground-based gimbaled payloads, like x-ray and gamma-ray telescopes. They can also be used to scan free-flying spacecraft. Circular scans, linear scans, and raster scans can be generated. A pair of RUM devices generates the basic scan motion and an auxiliary control system using torque motors, control moment gyros, or reaction wheels keeps the scan centered on the target and produces some complementary motion for raster scanning. Previous analyses and simulation results show that this approach offers significant power savings compared to scanning only with the auxiliary control system, especially with large payloads and high scan frequencies. However, these claims have never been proven until now. This paper describes a laboratory experiment which tests the concept of scanning a gimbaled payload with RUM devices. A description of the experiment is given and test results that prove the concept are presented. The test results are compared with those from a computer simulation model of the experiment and the differences are discussed.

KSC00pp1730

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister waits at the base of the Rotating Service Structure (RSS) with the P6 integrated truss segment inside. The canister will be lifted up to the payload changeout room (PCR) where the P6 will be removed for transfer to Space Shuttle Endeavour’s payload bay. The PCR is the enclosed, environmentally controlled portion of the RSS that supports payload delivery at the pad and subsequent vertical installation in the orbiter payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC-00pp1730

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister waits at the base of the Rotating Service Structure (RSS) with the P6 integrated truss segment inside. The canister will be lifted up to the payload changeout room (PCR) where the P6 will be removed for transfer to Space Shuttle Endeavour’s payload bay. The PCR is the enclosed, environmentally controlled portion of the RSS that supports payload delivery at the pad and subsequent vertical installation in the orbiter payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

Space Transportation System/Spacelab accommodations

NASA Technical Reports Server (NTRS)

De Sanctis, C. E.

1978-01-01

A description is provided of the capabilities offered by the Spacelab design for doing research in space. The Spacelab flight vehicle consists of two basic elements including the habitable pressurized compartments and the unpressurized equipment mounting platforms. Spacelab services to payloads are considered, taking into account payload mass, electrical power and energy, heat rejection for Spacelab and payload, aspects of Spacelab data handling, and the extended flight capability. Attention is also given to the Spacelab structure, crew station and habitability, the electrical power distribution subsystem, the command and data management subsystem, the experiment computer operating system, the environmental control subsystem, the experiment vent assembly, the common payload support equipment, the instrument pointing subsystem, and details concerning the utilization of Spacelab.

Nano Entry System for CubeSat-Class Payloads Project (Nano-ADEPT)

NASA Technical Reports Server (NTRS)

Smith, Brandon Patrick

2014-01-01

This project is developing a mechanically deployed system through a mission application study, deployment/ejection testing, and wind tunnel testing. Adaptable Deployable Entry and Placement Technology (ADEPT) has been under development at NASA since 2011. Nano-ADEPT is the application of this revolutionary entry technology for small spacecraft. The unique capability of ADEPT for small science payloads comes from its ability to stow within a slender volume and deploy passively to achieve a mass-efficient drag surface with a high heat rate capability. Near-term applications for this technology include return of small science payloads or CubeSat technology from Low Earth Orbit (LEO) and delivery of secondary payloads to the surface of Mars.

Neural controller for adaptive movements with unforeseen payloads.

PubMed

Kuperstein, M; Wang, J

1990-01-01

A theory and computer simulation of a neural controller that learns to move and position a link carrying an unforeseen payload accurately are presented. The neural controller learns adaptive dynamic control from its own experience. It does not use information about link mass, link length, or direction of gravity, and it uses only indirect uncalibrated information about payload and actuator limits. Its average positioning accuracy across a large range of payloads after learning is 3% of the positioning range. This neural controller can be used as a basis for coordinating any number of sensory inputs with limbs of any number of joints. The feedforward nature of control allows parallel implementation in real time across multiple joints.

An on-board processing satellite payload for European mobile communications

NASA Astrophysics Data System (ADS)

Evans, B. G.; Casewell, I. E.; Craig, A. D.

1987-06-01

An examination of the use of satellite on-board processing (OBP) for land mobile applications shows the feasibility of designing an OBP payload to satisfy the functional requirements of the land mobile system projected for the 1990s. Following a discussion of the proposed land mobile system, advantages of OBP over conventional transport payloads are considered. The use of digital signal processing techniques is shown to provide a solution for the merging of the routing and transmultiplexing functions into a single element, and such techniques are ideally suited to space applications. It is suggested that the projected power, mass, and size estimates are compatible with the payload capacity of one of the large Olympus satellites.

Design of High Frequency Pulse Tube Cryocooler for Onboard Space Applications

NASA Astrophysics Data System (ADS)

Srikanth, Thota; Padmanabhan; Gurudath, C. S.; Amrit, A.; Basavaraj, S.; Dinesh, K.

2017-02-01

To meet the growing demands of on-board applications such as cooling meteorological payloads and the satellite operational constraints like power, lower mass, reduced size and redundancy; a Pulse Tube Cryocooler (PTC) is designed by arriving at an operating frequency of 100 Hz and Helium gas pressure of 35 bar based on insights obtained from combination of phasor diagram, pulse tube and regenerator geometries with overall system mass of ≤ 2.0 kg. High frequency operation would allow reducing the size and mass of pressure wave modulator for a given input power. High Frequency also helps in reducing the volume of regenerator for a given cooling power, which increases the power density and leads to faster cool down. A component level modelling of the regenerator for optimising length and diameter for maximum Coefficient of Performance (COP) is carried out using REGEN3.3. The overall system level modelling of PTC is carried out using 1-D software SAGE. The cold end mass flow rate of the optimised regenerator is taken as reference for the system modelling. The performance achieved in REGEN3.3 is 2.15 W of net heat lift against the performance of 1.02 W of net heat lift at 80 K in SAGE.

Solar-powered unmanned aerial vehicles

DOE Office of Scientific and Technical Information (OSTI.GOV)

Reinhardt, K.C.; Lamp, T.R.; Geis, J.W.

1996-12-31

An analysis was performed to determine the impact of various power system components and mission requirements on the size of solar-powered high altitude long endurance (HALE)-type aircraft. The HALE unmanned aerial vehicle (UAV) has good potential for use in many military and civil applications. The primary power system components considered in this study were photovoltaic (PV) modules for power generation and regenerative fuel cells for energy storage. The impact of relevant component performance on UAV size and capability were considered; including PV module efficiency and mass, power electronics efficiency, and fuel cell specific energy. Mission parameters such as time ofmore » year, flight altitude, flight latitude, and payload mass and power were also varied to determine impact on UAV size. The aircraft analysis method used determines the required aircraft wing aspect ratio, wing area, and total mass based on maximum endurance or minimum required power calculations. The results indicate that the capacity of the energy storage system employed, fuel cells in this analysis, greatly impacts aircraft size, whereas the impact of PV module efficiency and mass is much less important. It was concluded that an energy storage specific energy (total system) of 250--500 Whr/kg is required to enable most useful missions, and that PV cells with efficiencies greater than {approximately} 12% are suitable for use.« less

The MICROSCOPE inertial sensor: qualification status

NASA Astrophysics Data System (ADS)

Santos Rodrigues, Manuel; Touboul, Pierre; Liorzou, Francoise; Bodoville, Guillaume

The payload of the MICROSCOPE space mission embarks two pairs of test-masses, made of Platinum Rhodium alloy and Titanium alloy, that are used to perform the test of the Uni-versality of free fall, i.e. of the Equivalence Principle (EP). These cylindrical test-masses are at the core of the inertial sensors used to perform the full drag-free and attitude control of the satellite. Based on electrostatic space accelerometers developed in ONERA, the payload has been designed with challenging technologies for the electronics and for the sensor core. Following a very specific development plan, the payload is currently in the qualification phase, by being integrated after a long period of challenging accurate production and metrology. The results obtained for the driving components of the expected performance will be addressed. In particular, the micrometric metrology of the instrument core, made in gold coated silica, will be presented: the specific ultra-sonic machining processes, optimized for this production, indeed exhibit a few micrometers accuracy. Similar accuracy is obtained for the geometry of the test-masses, produced and controlled in collaboration with the PTB, Physikalisch-Technische Bundesanstalt. This accurate geometry and the specific selected shape is mandatory to balance the mass moment of inertia for gravity gradient rejection and to highly decouple the instru-ment measurement axes. The first results of the flight model electronics will be also presented demonstrating micro-volt low noise and weak thermal sensitivity in good agreement with the requirements. At last, the development status of the payload will be mentioned insisting on the coming milestones.

Simulating flight boundary conditions for orbiter payload modal survey

NASA Technical Reports Server (NTRS)

Chung, Y. T.; Sernaker, M. L.; Peebles, J. H.

1993-01-01

An approach to simulate the characteristics of the payload/orbiter interfaces for the payload modal survey was developed. The flexure designed for this approach is required to provide adequate stiffness separation in the free and constrained interface degrees of freedom to closely resemble the flight boundary condition. Payloads will behave linearly and demonstrate similar modal effective mass distribution and load path as the flight if the flexure fixture is used for the payload modal survey. The potential non-linearities caused by the trunnion slippage during the conventional fixed base modal survey may be eliminated. Consequently, the effort to correlate the test and analysis models can be significantly reduced. An example is given to illustrate the selection and the sensitivity of the flexure stiffness. The advantages of using flexure fixtures for the modal survey and for the analytical model verification are also demonstrated.

Potential of extended airbreathing operation of a two-stage launch vehicle by scramjet propulsion

NASA Astrophysics Data System (ADS)

Schoettle, U. M.; Hillesheimer, M.; Rahn, M.

This paper examines the application of scramjet propulsion to extend the ramjet operation of an airbreathing two-stage launch designed for horizontal takeoff and landing. Performance comparisons are made for two alternative propulsion concepts. The mission performance predictions presented are obtained from a multistep optimization procedure employing both trajectory optimization and vehicle design steps to achieve maximum payload capabilities. The simulation results are shown to offer an attractive payload advantage of the scramjet variant over the ramjet powered vehicle.

Ares V Launch Capability Enables Future Space Telescopes

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2007-01-01

NASA's Ares V cargo launch vehicle offers the potential to completely change the paradigm of future space science mission architectures. A major finding of the NASA Advanced Telescope and Observatory Capability Roadmap Study was that current launch vehicle mass and volume constraints severely limit future space science missions. And thus, that significant technology development is required to package increasingly larger collecting apertures into existing launch shrouds. The Ares V greatly relaxes these constraints. For example, while a Delta IV has the ability to launch approximate a 4.5 meter diameter payload with a mass of 13,000 kg to L2, the Ares V is projected to have the ability to launch an 8 to 12 meter diameter payload with a mass of 60,000 kg to L2 and 130,000 kg to Low Earth Orbit. This paper summarizes the Ares V payload launch capability and introduces how it might enable new classes of future space telescopes such as 6 to 8 meter class monolithic primary mirror observatories, 15 meter class segmented telescopes, 6 to 8 meter class x-ray telescopes or high-energy particle calorimeters.

Ascent performance feasibility for next-generation spacecraft

NASA Astrophysics Data System (ADS)

Mancuso, Salvatore Massimo

This thesis deals with the optimization of the ascent trajectories for single-stage suborbital (SSSO), single-stage-to-orbit (SSTO), and two-stage-to-orbit (TSTO) rocket-powered spacecraft. The maximum payload weight problem has been solved using the sequential gradient-restoration algorithm. For the TSTO case, some modifications to the original version of the algorithm have been necessary in order to deal with discontinuities due to staging and the fact that the functional being minimized depends on interface conditions. The optimization problem is studied for different values of the initial thrust-to-weight ratio in the range 1.3 to 1.6, engine specific impulse in the range 400 to 500 sec, and spacecraft structural factor in the range 0.08 to 0.12. For the TSTO configuration, two subproblems are studied: uniform structural factor between stages and nonuniform structural factor between stages. Due to the regular behavior of the results obtained, engineering approximations have been developed which connect the maximum payload weight to the engine specific impulse and spacecraft structural factor; in turn, this leads to useful design considerations. Also, performance sensitivity to the scale of the aerodynamic drag is studied, and it is shown that its effect on payload weight is relatively small, even for drag changes approaching ± 50%. The main conclusions are that: the design of a SSSO configuration appears to be feasible; the design of a SSTO configuration might be comfortably feasible, marginally feasible, or unfeasible, depending on the parameter values assumed; the design of a TSTO configuration is not only feasible, but its payload appears to be considerably larger than that of a SSTO configuration. Improvements in engine specific impulse and spacecraft structural factor are desirable and crucial for SSTO feasibility; indeed, it appears that aerodynamic improvements do not yield significant improvements in payload weight.

The space shuttle payload planning working groups: Volume 9: Materials processing and space manufacturing

NASA Technical Reports Server (NTRS)

1973-01-01

The findings and recommendations of the Materials Processing and Space Manufacturing group of the space shuttle payload planning activity are presented. The effects of weightlessness on the levitation processes, mixture stability, and control over heat and mass transport in fluids are considered for investigation. The research and development projects include: (1) metallurgical processes, (2) electronic materials, (3) biological applications, and (4)nonmetallic materials and processes. Additional recommendations are provided concerning the allocation of payload space, acceptance of experiments for flight, flight qualification, and private use of the space shuttle.

Results of Evaluation of Solar Thermal Propulsion

NASA Technical Reports Server (NTRS)

Woodcock, Gordon; Byers, Dave

2003-01-01

The solar thermal propulsion evaluation reported here relied on prior research for all information on solar thermal propulsion technology and performance. Sources included personal contacts with experts in the field in addition to published reports and papers. Mission performance models were created based on this information in order to estimate performance and mass characteristics of solar thermal propulsion systems. Mission analysis was performed for a set of reference missions to assess the capabilities and benefits of solar thermal propulsion in comparison with alternative in-space propulsion systems such as chemical and electric propulsion. Mission analysis included estimation of delta V requirements as well as payload capabilities for a range of missions. Launch requirements and costs, and integration into launch vehicles, were also considered. The mission set included representative robotic scientific missions, and potential future NASA human missions beyond low Earth orbit. Commercial communications satellite delivery missions were also included, because if STP technology were selected for that application, frequent use is implied and this would help amortize costs for technology advancement and systems development. A C3 Topper mission was defined, calling for a relatively small STP. The application is to augment the launch energy (C3) available from launch vehicles with their built-in upper stages. Payload masses were obtained from references where available. The communications satellite masses represent the range of payload capabilities for the Delta IV Medium and/or Atlas launch vehicle family. Results indicated that STP could improve payload capability over current systems, but that this advantage cannot be realized except in a few cases because of payload fairing volume limitations on current launch vehicles. It was also found that acquiring a more capable (existing) launch vehicle, rather than adding an STP stage, is the most economical in most cases.

KSC-00pp1732

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, is lifted toward the payload changeout room (PCR). The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure that supports payload delivery at the pad and subsequent vertical installation in the orbiter payload bay. Attached to the canister are the red umbilical lines that maintain the controlled environment inside. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC00pp1732

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, is lifted toward the payload changeout room (PCR). The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure that supports payload delivery at the pad and subsequent vertical installation in the orbiter payload bay. Attached to the canister are the red umbilical lines that maintain the controlled environment inside. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

First Spacelab flight - A status report of the joint ESA/NASA mission

NASA Technical Reports Server (NTRS)

Craft, H. G., Jr.; Sanfourche, J.-P.

1978-01-01

A general overview of the first Spacelab flight is presented and a table is given listing the payload composition. An accommodation study is presented with emphasis on the configuration, mass status, timeline, and experiment interface specifications. Also considered are flight and ground operations, safety factors, and payload specialists training for the first flight.

SLS Payload Transportation Beyond LEO

NASA Technical Reports Server (NTRS)

Creech, S. D.; Baker, J. D.; Jackman, A. L.; Vane, G.

2017-01-01

NASA has successfully completed the Critical Design Review (CDR) of the heavy lift Space Launch System (SLS) and is working towards the first flight of the vehicle in 2018. SLS will begin flying crewed missions with an Orion capsule to the lunar vicinity every year after the first 2 flights starting in the early 2020's. As early as 2021, in addition to delivering an Orion capsule to a cislunar destination, SLS will also deliver ancillary payload, termed "Co-manifested Payload (CPL)", with a mass of at least 5.5 mT and volume up to 280 m3 simultaneously to that same destination. Later SLS flights have a goal of delivering as much as 10 mT of CPL to cislunar destinations. In addition to cislunar destinations, SLS flights may deliver non-crewed, science-driven missions with Primary Payload (PPL) to more distant destinations. SLS PPL missions will utilize a unique payload fairing offering payload volume (ranging from 320 m3 to 540 m3) that greatly exceeds the largest existing Expendable Launch Vehicle (ELV) fairing available. The Characteristic Energy (C3) offered by the SLS system will generate opportunities to deliver up to 40 mT to cislunar space, and deliver double PPL mass or de-crease flight time by half for some outer planet destinations when compared to existing capabilities. For example, SLS flights may deliver the Europa Clipper to a Jovian destination in under 3 years by the mid 2020's, compared to the 7+ years cruise time required for current launch capabilities. This presentation will describe ground and flight accommodations, interfaces, resources, and performance planned to be made available to potential CPL and PPL science users of SLS. In addition, this presentation should promote a dialogue between vehicle developers, potential payload users, and funding sources in order to most efficiently evolve required SLS capabilities to meet diverse payload needs as they are identified over the next 35 years and beyond.

A Spherical Torus Nuclear Fusion Reactor Space Propulsion Vehicle Concept for Fast Interplanetary Travel

NASA Technical Reports Server (NTRS)

Williams, Craig H.; Borowski, Stanley K.; Dudzinski, Leonard A.; Juhasz, Albert J.

1998-01-01

A conceptual vehicle design enabling fast outer solar system travel was produced predicated on a small aspect ratio spherical torus nuclear fusion reactor. Initial requirements were for a human mission to Saturn with a greater than 5% payload mass fraction and a one way trip time of less than one year. Analysis revealed that the vehicle could deliver a 108 mt crew habitat payload to Saturn rendezvous in 235 days, with an initial mass in low Earth orbit of 2,941 mt. Engineering conceptual design, analysis, and assessment was performed on all ma or systems including payload, central truss, nuclear reactor (including divertor and fuel injector), power conversion (including turbine, compressor, alternator, radiator, recuperator, and conditioning), magnetic nozzle, neutral beam injector, tankage, start/re-start reactor and battery, refrigeration, communications, reaction control, and in-space operations. Detailed assessment was done on reactor operations, including plasma characteristics, power balance, power utilization, and component design.

Technology Overview and Assessment for Small-Scale EDL Systems

NASA Technical Reports Server (NTRS)

Heidrich, Casey R.; Smith, Brandon P.; Braun, Robert D.

2016-01-01

Motivated by missions to land large rovers and humans at Mars and other bodies, high-mass EDL technologies are a prevalent trend in the research community. In contrast, EDL systems for low-mass payloads have attracted less attention. Significant potential in science and discovery exists in small-scale EDL systems. Payloads acting secondary to a flagship mission are a currently under-utilzed resource. Before taking advantage of these opportunities, further developed of scaled EDL technologies is required. The key limitations identified in this study are compact decelerators and deformable impact systems. Current technologies may enable rough landing of small payloads, with moderate restrictions in packaging volume. Utilization of passive descent and landing stages will greatly increase the applicability of small systems, allowing for vehicles robust to entry environment uncertainties. These architectures will provide an efficient means of achieving science and support objectives while reducing cost and risk margins of a parent mission.

Space Launch System Trans Lunar Payload Delivery Capability

NASA Technical Reports Server (NTRS)

Jackman, A. L.; Smith, D. A.

2016-01-01

NASA Marshall Space Flight Center (MSFC) has successfully completed the Critical Design Review (CDR) of the heavy lift Space Launch System (SLS) and is working towards first flight of the vehicle in 2018. SLS will begin flying crewed missions with an Orion to a lunar vicinity every year after the first 2 flights starting in the early 2020's. So as early as 2021 these Orion flights will deliver ancillary payload, termed "Co-Manifested Payload", with a mass of at least 5.5 metric tons and volume up to 280 cubic meters to a cis-lunar destination. Later SLS flights have a goal of delivering as much as 10 metric tons to a cis-lunar destination. This presentation will describe the ground and flight accommodations, interfaces, and resources planned to be made available to Co-Manifested Payload providers as part of the SLS system. An additional intention is to promote a two-way dialogue between vehicle developers and potential payload users in order to most efficiently evolve required SLS capabilities to meet diverse payload requirements.

The payloads of Advanced Virgo: current status and upgrades

NASA Astrophysics Data System (ADS)

Naticchioni, L.; Virgo Collaboration

2018-02-01

The development and integration of new detector payloads has been an important part of the Advanced Virgo (AdV) project, the major upgrade of the Virgo interferometric detector of Gravitational Waves, aiming to increase the detector sensitivity by one order of magnitude. During the integration phase of the new AdV payloads with monolithic suspension of mirrors we experienced systematic suspension failures later identified as caused by dust contamination of the vacuum system. In order to not postpone the detector commissioning, making possible to join the LIGO O2 observation run, the Collaboration decided to proceed with the integration of the payloads relying on steel wire suspensions for all the mirrors. In this proceeding the status of the currently integrated payloads is reported, including their angular control characterization and the Q-factor measurements for test mass steel wire suspensions. The payload upgrade for the re-integration of monolithic suspensions after the O2 run is reported in the last section.

Applications of a new mass-driver concept

NASA Technical Reports Server (NTRS)

Oneill, G. K.

1981-01-01

A description of the operating principles and requirements of a novel mass-driver concept is presented. The design obtains acceleration of payload bucket coils by means of transverse focussing from strong, off-axis restoring forces that are produced by drive coils operating in a 'pull-only' mode. The concept offers the unprecedented possibility of operating high-performance mass-drivers entirely within the limitations of existing commercial switching devices, such as silicon-controlled rectifiers, spark gaps, vacuum-triggered arcs or vacuum mechanical switches. Representative applications of the concept described are: (1) a large-diameter magnetic lunar launcher for payloads having autonomous maneuvering; (2) an intermediate-diameter launcher with long operational life; and (3) a reaction engine for orbit transfer of large, massive objects.

Mass breakdown model of solar-photon sail shuttle: The case for Mars

NASA Astrophysics Data System (ADS)

Vulpetti, Giovanni; Circi, Christian

2016-02-01

The main aim of this paper is to set up a many-parameter model of mass breakdown to be applied to a reusable Earth-Mars-Earth solar-photon sail shuttle, and analyze the system behavior in two sub-problems: (1) the zero-payload shuttle, and (2) given the sailcraft sail loading and the gross payload mass, find the sail area of the shuttle. The solution to the subproblem-1 is of technological and programmatic importance. The general analysis of subproblem-2 is presented as a function of the sail side length, system mass, sail loading and thickness. In addition to the behaviors of the main system masses, useful information for future work on the sailcraft trajectory optimization is obtained via (a) a detailed mass model for the descent/ascent Martian Excursion Module, and (b) the fifty-fifty solution to the sailcraft sail loading breakdown equation. Of considerable importance is the evaluation of the minimum altitude for the rendezvous between the ascent rocket vehicle and the solar-photon sail propulsion module, a task performed via the Mars Climate Database 2014-2015. The analysis shows that such altitude is 300 km; below it, the atmospheric drag prevails over the solar-radiation thrust. By this value, an example of excursion module of 1500 kg in total mass is built, and the sailcraft sail loading and the return payload are calculated. Finally, the concept of launch opportunity-wide for a shuttle driven by solar-photon sail is introduced. The previous fifty-fifty solution may be a good initial guess for the trajectory optimization of this type of shuttle.

Commercially Hosted Government Payloads: Lessons from Recent Programs

NASA Technical Reports Server (NTRS)

Andraschko, Mark A.; Antol, Jeffrey; Horan, Stephen; Neil, Doreen

2011-01-01

In a commercially hosted operational mode, a scientific instrument or operational device is attached to a spacecraft but operates independently from the spacecraft s primary mission. Despite the expected benefits of this arrangement, there are few examples of hosted payload programs actually being executed by government organizations. The lack of hosted payload programs is largely driven by programmatic challenges, both real and perceived, rather than by technical challenges. Partly for these reasons, NASA has not sponsored a hosted payload program, in spite of the benefits and visible community interest in doing so. In the interest of increasing the use of hosted payloads across the space community, this paper seeks to alleviate concerns about hosted payloads by identifying these programmatic challenges and presenting ways in which they can be avoided or mitigated. Despite the challenges, several recent hosted payload programs have been successfully completed or are currently in progress. This paper presents an assessment of these programs, with a focus on acquisition, costs, schedules, risks, and other programmatic aspects. The hosted payloads included in this study are the Federal Aviation Administration's Wide Area Augmentation System (WAAS) payloads, United States Coast Guard's Automatic Identification System (AIS) demonstration payload, Department of Defense's IP Router In Space (IRIS) demonstration payload, the United States Air Force's Commercially Hosted Infrared Payload (CHIRP), and the Australian Defence Force's Ultra High Frequency (UHF) payload. General descriptions of each of these programs are presented along with issues that have been encountered and lessons learned from those experiences. A set of recommended approaches for future hosted payload programs is presented, with a focus on addressing risks or potential problem areas through smart and flexible contracting up front. This set of lessons and recommendations is broadly applicable to future hosted payload programs, whether they are technology demonstrations, communications systems, or operational sensors. Additionally, we present a basic cost model for commercial access to space for hosted payloads as a function of payload mass

Adaptive output-based command shaping for sway control of a 3D overhead crane with payload hoisting and wind disturbance

NASA Astrophysics Data System (ADS)

Abdullahi, Auwalu M.; Mohamed, Z.; Selamat, H.; Pota, Hemanshu R.; Zainal Abidin, M. S.; Ismail, F. S.; Haruna, A.

2018-01-01

Payload hoisting and wind disturbance during crane operations are among the challenging factors that affect a payload sway and thus, affect the crane's performance. This paper proposes a new online adaptive output-based command shaping (AOCS) technique for an effective payload sway reduction of an overhead crane under the influence of those effects. This technique enhances the previously developed output-based command shaping (OCS) which was effective only for a fixed system and without external disturbances. Unlike the conventional input shaping design technique which requires the system's natural frequency and damping ratio, the proposed technique is designed by using the output signal and thus, an online adaptive algorithm can be formulated. To test the effectiveness of the AOCS, experiments are carried out using a laboratory overhead crane with a payload hoisting in the presence of wind, and with different payloads. The superiority of the method is confirmed by 82% and 29% reductions in the overall sway and the maximum transient sway respectively, when compared to the OCS, and two robust input shapers namely Zero Vibration Derivative-Derivative and Extra-Insensitive shapers. Furthermore, the method demonstrates a uniform crane's performance under all conditions. It is envisaged that the proposed method can be very useful in designing an effective controller for a crane system with an unknown payload and under the influence of external disturbances.

Operation of a swept Langmuir probe on a sounding rocket

NASA Astrophysics Data System (ADS)

Robertson, S. H.; Dickson, S.; Friedrich, M.; Sternovsky, Z.

2012-12-01

A swept cylindrical Langmuir probe was operated on two sounding rockets from ~ 60-120 km for the purpose of determining both the ambient electron density and the payload potential relative to the ambient plasma. The rockets were part of the CHAMPS (CHarge And mass of Meteoritic smoke ParticleS) rocket campaign and carried mass analyzers and various plasma probes to study charged meteoritic dust in the mesopause region. The payload potential is an important parameter for data interpretation. The rockets were launched in October of 2011 from Andøya Rocket Range, Norway. The launches were a few days apart with one taking place during the day and the other at night. The swept Langmuir probe data provided a current-voltage characteristic that had a distinct "knee" indicating the onset of electron collection; the probe voltage at this "knee" corresponds to the ambient plasma potential. The data indicate a payload potential of about -2 V to -1 V for both launches. The payload potential becomes less negative for altitudes above 80 km on the day launch due to photoemission. The probe current-voltage data are also compared with ion and electron density measurements from ion probes and Faraday rotation antennas, respectively. The data from the various instruments are in general agreement. Further consideration of the Langmuir probe performance shows that if the probe had been operated with feedback control to continuously collect electrons with a current of order 1 microamp, the probe potential would be an accurate, continuous indicator of the payload potential without the need for sweeping which could periodically alter the payload potential.

Maximizing Launch Vehicle and Payload Design Via Early Communications

NASA Technical Reports Server (NTRS)

Morris, Bruce

2010-01-01

The United States? current fleet of launch vehicles is largely derived from decades-old designs originally made for payloads that no longer exist. They were built primarily for national security or human exploration missions. Today that fleet can be divided roughly into small-, medium-, and large-payload classes based on mass and volume capability. But no vehicle in the U.S. fleet is designed to accommodate modern payloads. It is usually the payloads that must accommodate the capabilities of the launch vehicles. This is perhaps most true of science payloads. It was this paradigm that the organizers of two weekend workshops in 2008 at NASA's Ames Research Center sought to alter. The workshops brought together designers of NASA's Ares V cargo launch vehicle (CLV) with scientists and payload designers in the astronomy and planetary sciences communities. Ares V was still in a pre-concept development phase as part of NASA?s Constellation Program for exploration beyond low Earth orbit (LEO). The space science community was early in a Decadal Survey that would determine future priorities for research areas, observations, and notional missions to make those observations. The primary purpose of the meetings in April and August of 2008, including the novel format, was to bring vehicle designers together with space scientists to discuss the feasibility of using a heavy lift capability to launch large observatories and explore the Solar System. A key question put to the science community was whether this heavy lift capability enabled or enhanced breakthrough science. The meetings also raised the question of whether some trade-off between mass/volume and technical complexity existed that could reduce technical and programmatic risk. By engaging the scientific community early in the vehicle design process, vehicle engineers sought to better understand potential limitations and requirements that could be added to the Ares V from the mission planning community. From the vehicle standpoint, while the human exploration mission could not be compromised to accommodate other payloads, the design might otherwise be tailored to not exclude other payload requirements. This paper summarizes the findings of the workshops and discusses the benefits of bringing together the vehicle design and science communities early in their concept phases

A concept of highly maneuverable experimental space (HIMES) vehicle

NASA Astrophysics Data System (ADS)

Nagatomo, M.; Naruo, Y.; Inatani, Y.

1985-10-01

The development of a highly maneuverable experimental space (HIMES) vehicle is proposed. This reusable sounding rocket is to be propelled by a liquid hydrogen/LOX engine, and have a maximum payload mass of 500 kg at an altitude of 300 km. The main subsystems of HIMES, the fuselage and wing structure, propulsion, and navigation, guidance, and control system, are described and a diagram is provided. The operational features of HIMES are defined by three mission models. In the first model the vehicle is used as a stable platform of low velocity relative to the environment; model two represents the suborbital flight of sounding rockets, and model three is used for orbital reentry experiments and the testing of a new system of winged space vehicles. Typical mission profiles for the three models are presented. A cost estimation of the HIMES vehicle is given.

An optical approach to proximity-operations communications for Space Station Freedom

NASA Technical Reports Server (NTRS)

Marshalek, Robert G.

1991-01-01

An optical communications system is described that supports bi-directional interconnections between Space Station Freedom (SSF) and a host of attached and co-orbiting platforms. These proximity-operations (Prox-Ops) platforms are categorized by their maximum distance from SSF, with several remaining inside 1-km range and several extending out to 37-km and 2000-km ranges in the initial and growth phases, respectively. Two distinct Prox-Ops optical terminals are described. A 1-cm-aperture system is used on the short-range platforms to reduce payload mass, and a 10-cm-aperture system is used on the long-range platforms and on SSF to support the optical link budgets. The system supports up to four simultaneous user links, by assigning wavelengths to the various platforms and by using separate SSF terminals for each link.

Radiation energy receiver for laser and solar propulsion systems

NASA Technical Reports Server (NTRS)

Rault, D. F. G.; Hertzberg, A.

1983-01-01

The concept of remotely heating a rocket propellant with a high intensity radiant energy flux is especially attractive due to its high specific impulse and large payload mass capabilities. In this paper, a radiation receiver-thruster which is especially suited to the particular thermodynamic and spectral characteristics of highly concentrated solar energy is proposed. In this receiver, radiant energy is volumetrically absorbed within a hydrogen gas seeded with alkali metal vapors. The alkali atoms and molecules absorb the radiant flux and, subsequently, transfer their internal excitation to hydrogen molecules through collisional quenching. It is shown that such a radiation receiver would outperform a blackbody cavity type receiver in both efficiency and maximum operating temperatures. A solar rocket equipped with such a receiver-thruster would deliver thrusts of several hundred newtons at a specific impulse of 1000 seconds.

Probabilistic Sensitivity Analysis for Launch Vehicles with Varying Payloads and Adapters for Structural Dynamics and Loads

NASA Technical Reports Server (NTRS)

McGhee, David S.; Peck, Jeff A.; McDonald, Emmett J.

2012-01-01

This paper examines Probabilistic Sensitivity Analysis (PSA) methods and tools in an effort to understand their utility in vehicle loads and dynamic analysis. Specifically, this study addresses how these methods may be used to establish limits on payload mass and cg location and requirements on adaptor stiffnesses while maintaining vehicle loads and frequencies within established bounds. To this end, PSA methods and tools are applied to a realistic, but manageable, integrated launch vehicle analysis where payload and payload adaptor parameters are modeled as random variables. This analysis is used to study both Regional Response PSA (RRPSA) and Global Response PSA (GRPSA) methods, with a primary focus on sampling based techniques. For contrast, some MPP based approaches are also examined.

The cyclotron energization through auroral wave experiments (CENTAUR 2B)

NASA Technical Reports Server (NTRS)

Winningham, J. D.

1992-01-01

The CENTAUR 2B mission, a dual payload program, is in many aspects the same as the previous missions from Cape Perry and Norway in 1985. It was planned that these payloads would be launched from Andoya, Norway, Nov. 1989 from the Universal II launcher. The payloads are identical, but would have been launched at different azimuths as far north and as far west as possible. Particle experiments include the angular resolving energy analyzer (AREA), the fast ion mass spectrometer (FIMS), the spectrographic particle images (SPI), and finally, the differential ion flux probe (DIFP). SwRI was responsible for the scientific payload, which includes the power supplies, the power supply interfacing, the manipulating of the data from the instruments to format it for the telemetry system, all mechanical structure and restraint mechanisms, and the payload subskin. The status of the various components of this program is given.

International Space Station-Based Electromagnetic Launcher for Space Science Payloads

NASA Technical Reports Server (NTRS)

Jones, Ross M.

2013-01-01

A method was developed of lowering the cost of planetary exploration missions by using an electromagnetic propulsion/launcher, rather than a chemical-fueled rocket for propulsion. An electromagnetic launcher (EML) based at the International Space Station (ISS) would be used to launch small science payloads to the Moon and near Earth asteroids (NEAs) for the science and exploration missions. An ISS-based electromagnetic launcher could also inject science payloads into orbits around the Earth and perhaps to Mars. The EML would replace rocket technology for certain missions. The EML is a high-energy system that uses electricity rather than propellant to accelerate payloads to high velocities. The most common type of EML is the rail gun. Other types are possible, e.g., a coil gun, also known as a Gauss gun or mass driver. The EML could also "drop" science payloads into the Earth's upper

Assessment of the Influence of the RaD-X Balloon Payload on the Onboard Radiation Detectors

NASA Technical Reports Server (NTRS)

Gronoff, Guilluame; Mertens, Christopher J.; Norman, Ryan B.; Straume, Tore; Lusby, Terry C.

2016-01-01

The NASA Radiation Dosimetry Experiment (RaD-X) stratospheric balloon flight mission, launched on 25 September 2015, provided dosimetric measurements above the Pfotzer maximum. The goal of taking these measurements is to improve aviation radiation models by providing a characterization of cosmic ray primaries, which are the source of radiation exposure at aviation altitudes. The RaD-X science payload consists of four instruments. The main science instrument is a tissue-equivalent proportional counter (TEPC). The other instruments consisted of three solid state silicon dosimeters: Liulin, Teledyne total ionizing dose (TID) and RaySure detectors. The instruments were housed in an aluminum structure protected by a foam cover. The structure partially shielded the detectors from cosmic rays but also created secondary particles, modifying the ambient radiation environment observed by the instruments. Therefore, it is necessary to account for the influence of the payload structure on the measured doses. In this paper, we present the results of modeling the effect of the balloon payload on the radiation detector measurements using a Geant-4 (GEometry ANd Tracking) application. Payload structure correction factors derived for the TEPC, Liulin, and TID instruments are provided as a function of altitude. Overall, the payload corrections are no more than a 7% effect on the radiation environment measurements.

STS-39 AFP-675 and STP-1 MPESS in OV-103's payload bay (PLB)

NASA Image and Video Library

1991-05-06

STS039-10-019 (28 April-6 May 1991) --- This 35mm frame, taken from inside the crew cabin, shows some of the cargo in Discovery's payload bay. Seen are the tops of canisters on the STP-1 payload, configured on the STS 39 Hitchhiker carrier; and the Air Force Program (AFP) 675 package. AFP-675 consists of the Cryogenic Infrared Radiance Instrumentation for Shuttle (CIRRIS)-1A; Far Ultraviolet Camera (FAR-UV) Experiment; Horizon Ultraviolet Program (HUP); Quadruple Ion Neutral Mass Spectrometer (QINMS); and the Uniformly Redundant Array (URA).

Mars 2001 Lander Mission: Measurement Synergy through Coordinated Operations Planning and Implementation

NASA Astrophysics Data System (ADS)

Arvidson, R.; Bell, J. F., III; Kaplan, D.; Marshall, J.; Mishkin, A.; Saunders, S.; Smith, P.; Squyres, S.

1999-03-01

The Science Operations Working Group, Mars 2001 Mission, has developed coordinated plans for scientific observations that treat the instruments as an integrated payload. This approach ensures maximum return of scientific information.

Results from Two Low Mass Cosmic Ray Experiments Flown on the HASP Platform

NASA Astrophysics Data System (ADS)

Fontenot, R. S.; Hollerman, W. A.; Tittsworth, M.; Fountain, W.; Christl, M.; Thibodaux, C.; Broussard, B. M.

2009-03-01

The High Altitude Student Payload (HASP) program is designed to carry twelve student experiments to an altitude of about 123,000 feet (˜37 km). In 2006, students participated in the first HASP launch to measure cosmic ray intensities using traditional film and absorbers. This 10 kg payload flew from Fort Sumner, New Mexico in early September 2006 and was a great success. In 2007, students participated in the second HASP flight to measure the cosmic ray intensity and flux using a traditional film and absorber stack with five layers of optically stimulated luminescent (OSL) dosimeters. Results from both payloads showed that the cosmic ray flux decreases as a function of payload depth. As the cosmic rays go through the stack, they deposit their energy in the payload material. Determining cosmic ray flux is a tedious task. It involves digitizing the film and determining the real cosmic ray density. For the first HASP payload, students used a program known as GlobalLab to count particles. For the second payload, the students decided to use a combination of the GREYCStoration image regularization algorithm, an embossing filter, and a depth-merging filter to reconstruct the paths of the cosmic rays.

Technical Report: Unmanned Helicopter Solution for Survey-Grade Lidar and Hyperspectral Mapping

NASA Astrophysics Data System (ADS)

Kaňuk, Ján; Gallay, Michal; Eck, Christoph; Zgraggen, Carlo; Dvorný, Eduard

2018-05-01

Recent development of light-weight unmanned airborne vehicles (UAV) and miniaturization of sensors provide new possibilities for remote sensing and high-resolution mapping. Mini-UAV platforms are emerging, but powerful UAV platforms of higher payload capacity are required to carry the sensors for survey-grade mapping. In this paper, we demonstrate a technological solution and application of two different payloads for highly accurate and detailed mapping. The unmanned airborne system (UAS) comprises a Scout B1-100 autonomously operating UAV helicopter powered by a gasoline two-stroke engine with maximum take-off weight of 75 kg. The UAV allows for integrating of up to 18 kg of a customized payload. Our technological solution comprises two types of payload completely independent of the platform. The first payload contains a VUX-1 laser scanner (Riegl, Austria) and a Sony A6000 E-Mount photo camera. The second payload integrates a hyperspectral push-broom scanner AISA Kestrel 10 (Specim, Finland). The two payloads need to be alternated if mapping with both is required. Both payloads include an inertial navigation system xNAV550 (Oxford Technical Solutions Ltd., United Kingdom), a separate data link, and a power supply unit. Such a constellation allowed for achieving high accuracy of the flight line post-processing in two test missions. The standard deviation was 0.02 m (XY) and 0.025 m (Z), respectively. The intended application of the UAS was for high-resolution mapping and monitoring of landscape dynamics (landslides, erosion, flooding, or crops growth). The legal regulations for such UAV applications in Switzerland and Slovakia are also discussed.

P-MASS and P-GBA: Two new hardware developments for growing plants in space

NASA Technical Reports Server (NTRS)

Hoehn, Alexander; Luttges, Marvin W.; Robinson, Michael C.; Stodieck, Louis S.; Kliss, Mark H.

1994-01-01

Plant growth, and especially plant performance experiments in microgravity are limited by the currently available plant growth facilities (low light levels, inadequate nutrient delivery and atmosphere conditioning systems, insufficient science instrumentation, infrequent flight opportunities). In addition, mission durations of 10 to 14 days aboard the NSTS Space Shuttle allow for only brief periods of microgravity exposure with respect to the life cycle of a plant. Based on seed germination experiments, using the Generic BioProcessing Apparatus hardware (GBA), two new payloads have been designed specifically for plant growth. These payloads provide new opportunities for plant gravitational and space biology research and emphasize the investigation of plant performance (photosynthesis, biomass accumulations) in microgravity. The Plant-Module for Autonomous Space Support (P-MASS) was designed to utilize microgravity exposure times in excess of 30 days on the first flight of the recoverable COMET satellite (Commercial Experiment Transporter). The Plant-Generic Bioprocessing Apparatus (P-GBA), is designed for the National Space Transportation System (NSTS) Space Shuttle middeck and the SPACEHAB environment. The P-GBA is an evolution from the GBA hardware and P-MASS (plant chamber and instrumentation). The available light levels of both payloads more than double currently available capabilities.

Error tracking control for underactuated overhead cranes against arbitrary initial payload swing angles

NASA Astrophysics Data System (ADS)

Zhang, Menghua; Ma, Xin; Rong, Xuewen; Tian, Xincheng; Li, Yibin

2017-02-01

This paper exploits an error tracking control method for overhead crane systems for which the error trajectories for the trolley and the payload swing can be pre-specified. The proposed method does not require that the initial payload swing angle remains zero, whereas this requirement is usually assumed in conventional methods. The significant feature of the proposed method is its superior control performance as well as its strong robustness over different or uncertain rope lengths, payload masses, desired positions, initial payload swing angles, and external disturbances. Owing to the same attenuation behavior, the desired error trajectory for the trolley for each traveling distance is not needed to be reset, which is easy to implement in practical applications. By converting the error tracking overhead crane dynamics to the objective system, we obtain the error tracking control law for arbitrary initial payload swing angles. Lyapunov techniques and LaSalle's invariance theorem are utilized to prove the convergence and stability of the closed-loop system. Simulation and experimental results are illustrated to validate the superior performance of the proposed error tracking control method.

KSC-00pp1736

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- The payload transport canister, with the P6 integrated truss segment inside, is close to the payload changeout room on the Rotating Service Structure (RSS) at left. The PCR is the enclosed, environmentally controlled portion of the RSS that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and one solid rocket booster showing behind it. When the RSS is closed around Endeavour, the P6 truss will be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC00pp1736

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- The payload transport canister, with the P6 integrated truss segment inside, is close to the payload changeout room on the Rotating Service Structure (RSS) at left. The PCR is the enclosed, environmentally controlled portion of the RSS that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and one solid rocket booster showing behind it. When the RSS is closed around Endeavour, the P6 truss will be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

Automation of a Versatile Crane (the LSMS) for Lunar Outpost Construction, Maintenance and Inspection

NASA Technical Reports Server (NTRS)

Doggett, William R.; Roithmayr, Carlos M.; Dorsey, John T.; Jones, Thomas C.; Shen, Haijun; Seywald, Hans; King, Bruce D.; Mikulas, Martin M., Jr.

2009-01-01

Devices for lifting, translating and precisely placing payloads are critical for efficient Earth-based construction operations. Both recent and past studies have demonstrated that devices with similar functionality will be needed to support lunar outpost operations. Although several designs have been developed for Earth based applications, these devices lack unique design characteristics necessary for transport to and use on the harsh lunar surface. These design characteristics include: a) lightweight components, b) compact packaging for launch, c) automated deployment, d) simple in-field reconfiguration and repair, and e) support for tele-operated or automated operations. Also, because the cost to transport mass to the lunar surface is very high, the number of devices that can be dedicated to surface operations will be limited. Thus, in contrast to Earth-based construction, where many single-purpose devices dominate a construction site, a lunar outpost will require a limited number of versatile devices that provide operational benefit from initial construction through sustained operations. The first generation test-bed of a new high performance device, the Lunar Surface Manipulation System (LSMS) has been designed, built and field tested. The LSMS has many unique features resulting in a mass efficient solution to payload handling on the lunar surface. Typically, the LSMS device mass is estimated at approximately 3% of the mass of the heaviest payload lifted at the tip, or 1.8 % of the mass of the heaviest mass lifted at the elbow or mid-span of the boom for a high performance variant incorporating advanced structural components. Initial operational capabilities of the LSMS were successfully demonstrated during field tests at Moses Lake, Washington using a tele-operated approach. Joint angle sensors have been developed for the LSMS to improve operator situational awareness. These same sensors provide the necessary information to support fully automated operations, greatly expanding the operational versatility of the LSMS. This paper develops the equations describing the forward and inverse relation between LSMS joint angles and Cartesian coordinates of the LSMS tip. These equations allow a variety of schemes to be used to maneuver the LSMS to optimize the maneuver. One such scheme will be described in detail that eliminates undesirable swinging of the payload at the conclusion of a maneuver, even when the payload is suspended from a passive rigid link. The swinging is undesirable when performing precision maneuvers, such as aligning an object for mating or positioning a camera. Use of the equations described here enables automated control of the LSMS greatly improving its operational versatility.

Integration and Ruggedization of a Commercially Available Gas Chromatograph and Mass Spectrometer (GCMS) for the Resource Prospector Mission (RPM)

NASA Technical Reports Server (NTRS)

Loftin, Kathleen; Griffin, Timothy; Captain, Janine

2013-01-01

The Resource Prospector is a mission to prospect for lunar volatiles (primarily water) at one of the two lunar poles, as well as demonstrate In-Situ Resource Utilization (ISRU) on the Moon. The Resource Prospector consists of a lander, a rover, and a rover-borne scientific payload. The Regolith and Environment Science and Oxygen & Lunar Volatile Extraction (RESOLVE) payload, will be able to (1) locate near subsurface volatiles, (2) excavate and analyze samples of the volatile-bearing regolith, and (3) demonstrate the form, extractability and usefulness of the materials. The gas chromatograph mass spectrometer (GCMS) is the primary instrument in the RESOLVE instrumentation suite responsible for identification and quantification of the volatiles evolved from the lunar regolith. Specifically, this instrument must have: a low mass, a low power consumption, be able to perform fast analyses of samples ranging from less than one to greater than ninety nine percent water by mass, be autonomously controlled by the payload's software and avionics platform, and be able to operate in the harsh lunar environment. The RPM's short mission duration is the primary driver of the requirement for a very fast analysis time currently base lined at less than 2 minutes per sample. This presentation will discuss the requirements levied upon the GCMS design, lessons learned from a preliminary field demonstration deployment, the current design, and the path forward.

Plant-module for autonomous space-support (p-mass).

NASA Technical Reports Server (NTRS)

Luttges, M. W.; Stodieck, L.; Hoehn, A.

1994-01-01

A wide variety of technical and science questions arise when attempting to envision the long-term support of plants, algae and bacteria in space. Currently, spaceflight data remain elusive since there are no U.S. carriers for investigating either the germane technical or scientific issues. The first flight of the Commercial Experiment Transporter (COMET) will provide a nominal 30 day orbital opportunity to evlauate such issues. The P-Mass is a small payload that is designed to meet the mass (40 lbs), Volume (1.5cu. ft.), and power (120 W0 constraints of one of several COMET payloads while enabling flight evaluations of plants, algae and bacteria. Various biological sub-systems have been similarly evaluated. Through a variety of sensors coupled with color video, the P-Mass performance and the supported biological systems will be compared for terrestrial controls versus spaceflight materials. This small, low cost payload should return valuable regarding the requirements for hardware and biological systems needed to move toward biogenerative life support systems in space. In addition, it should be possible to accurately identify major unresolved difficulties that may arise in the long-term, this generic spaceflight capability should enable a variety of plant research programs focused on the use of microgravity to modulate and exploit plant products for commercial applications ranging from new agricultural products to pharmacological feedstocks and new controlled agricultural strategies.

Sounding rocket flight report: MUMP 9 and MUMP 10

NASA Technical Reports Server (NTRS)

Grassl, H. J.

1971-01-01

The results of the launching of two Marshall-University of Michigan Probes (MUMP 9 and MUMP 10), Nike-Tomahawk sounding rocket payloads, are summarized. The MUMP 9 paylaod included an omegatron mass analyzer, a molecular fluorescence densitometer, a mini-tilty filter, and a lunar position sensor. This complement of instruments permitted the determination of the molecular nitrogen density and temperature in the altitude range from approximately 143 to 297 km over Wallops Island, Virginia, during January 1971. The MUMP 10 payload included an omegatron mass analyzer, an electron temperature probe (Spencer, Brace, and Carignan, 1962), a cryogenic densitometer, and a solar position sensor. This complement of instruments permitted the determination of the molecular nitrogen density and temperature and the charged particle density and temperature in the altitude range from approximately 145 to 290 km over Wallops Island, Virginia, during the afternoon preceding the MUMP 9 launch in January 1971. A general description of the payload kinematics, orientation analysis, and the technique for the reduction and analysis of the data is given.

Sounding rocket flight report, MUMP 9 and MUMP 10

NASA Technical Reports Server (NTRS)

Grassl, H. J.

1971-01-01

The results of the launching of two-Marshall-University of Michigan Probes (MUMP 9 and MUMP 10), Nike-Tomahawk sounding rocket payloads, are summarized. The MUMP is similar to the thermosphere probe, an ejectable instrument package for studying the variability of the earth's atmospheric parameters. The MUMP 9 payload included an omegatron mass analyzer, a molecular fluorescence densitometer, a mini-tilty filter, and a lunar position sensor. This complement of instruments permitted the determination of the molecular nitrogen density and temperature in the altitude range from approximately 143 to 297 km over Wallops Island, Virginia, during January 1971. The MUMP 10 payload included an omegatron mass analyzer, an electron temperature probe, a cryogenic densitometer, and a solar position sensor. These instruments permitted the determination of the molecular nitrogen density and temperature and the charged particle density and temperature in the altitude range from approximately 145 to 290 km over Wallops Island during the afternoon preceding the MUMP 9 launch.

Probe design

NASA Technical Reports Server (NTRS)

Cowan, W.

1974-01-01

Outer planetary probe designs consider mission characteristics, structural configuration, delivery mode, scientific payload, environmental extremes, mass properties, and the launch vehicle and spacecraft interface.

Methods of and system for swing damping movement of suspended objects

DOEpatents

Jones, J.F.; Petterson, B.J.; Strip, D.R.

1991-03-05

A payload suspended from a gantry is swing damped in accordance with a control algorithm based on the periodic motion of the suspended mass or by servoing on the forces induced by the suspended mass. 13 figures.

The Cloud-Aerosol Transport System (CATS): A New Earth Science Capability for ISS (Invited)

NASA Astrophysics Data System (ADS)

McGill, M. J.; Yorks, J. E.; Scott, S.; Kupchock, A.; Selmer, P.

2013-12-01

The Cloud-Aerosol Transport System (CATS) is a lidar remote sensing instrument developed for deployment to the International Space Station (ISS). The CATS lidar will provide range-resolved profile measurements of atmospheric aerosol and cloud distributions and properties. The CATS instrument uses a high repetition rate laser operating at three wavelengths (1064, 532, and 355 nm) to derive properties of cloud/aerosol layers including: layer height, layer thickness, backscatter, optical depth, extinction, and depolarization-based discrimination of particle type. The CATS mission was designed to capitalize on the Space Station's unique orbit and facilities to continue existing Earth Science data records, to provide observational data for use in forecast models, and to demonstrate new technologies for use in future missions. The CATS payload will be installed on the Japanese Experiment Module - Exposed Facility (JEM-EF). The payload is designed to operate on-orbit for at least six months, and up to three years. The payload is completed and currently scheduled for a mid-2014 launch. The ISS and, in particular, the JEM-EF, is an exciting new platform for spaceborne Earth observations. The ability to leverage existing aircraft instrument designs coupled with the lower cost possible for ISS external attached payloads permits rapid and cost effective development of spaceborne sensors. The CATS payload is based on existing instrumentation built and operated on the high-altitude NASA ER-2 aircraft. The payload is housed in a 1.5 m x 1 m x 0.8 m volume that attaches to the JEM-EF. The allowed volume limits the maximum size for the collecting telescope to 60 cm diameter. Figure 1 shows a schematic layout of the CATS payload, with the primary instrument components identified. Figure 2 is a photo of the completed payload. CATS payload cut-away view. Completed CATS payload assembly.

Mars Sample Return Using Commercial Capabilities: Propulsive Entry, Descent, and Landing of a Capsule Form Vehicle

NASA Technical Reports Server (NTRS)

Gonzales, Andrew A.; Lemke, Lawrence G.; Huynh, Loc C.

2014-01-01

This paper describes a critical portion of the work that has been done at NASA, Ames Research Center regarding the use of the commercially developed Dragon capsule as a delivery vehicle for the elements of a high priority Mars Sample Return mission. The objective of the investigation was to determine entry and landed mass capabilities that cover anticipated mission conditions. The "Red Dragon", Mars configuration, uses supersonic retro-propulsion, with no required parachute system, to perform Entry, Descent, and Landing (EDL) maneuvers. The propulsive system proposed for use is the same system that will perform an abort, if necessary, for a human rated version of the Dragon capsule. Standard trajectory analysis tools are applied to publically available information about Dragon and other legacy capsule forms in order to perform the investigation. Trajectory simulation parameters include entry velocity, flight path angle, lift to drag Ratio (L/D), landing site elevation, atmosphere density, and total entry mass, in addition engineering assumptions for the performance of the propulsion system are stated. Mass estimates for major elements of the overall proposed architecture are coupled to this EDL analysis to close the overall architecture. Three synodic launch opportunities, beginning with the 2022 opportunity, define the arrival conditions. Results state the relations between the analysis parameters as well as sensitivities to those parameters. The EDL performance envelope includes landing altitudes between 0 and -4 km referenced to the Mars Orbiter Laser Altimeter datum as well as minimum and maximum atmosphere density. Total entry masses between 7 and 10 mt are considered with architecture closure occurring between 9.0 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the Thermal Protection System (TPS) currently in use for Dragon missions shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass and grow allowance for a Mars Ascent Vehicle (MAV), Earth Return Vehicle (ERV), and mission unique equipment. The useful payload supports an architecture that receives a sample from another surface asset and sends it directly back to Earth for recovery in a high Earth orbit. The work shows that emerging commercial capabilities as well as previously studied EDL methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that will also use propulsive EDL techniques

Microrover Nanokhod enhancing the scientific output of the ExoMars Lander

NASA Astrophysics Data System (ADS)

Klinkner, Sabine; Bernhardt, Bodo; Henkel, Hartmut; Rodionov, Daniel; Klingelhoefer, Goestar

The Nanokhod rover is a small and mobile exploration platform carrying out in-situ exploration by transporting and operating scientific instruments to interesting samples beyond the landing point. The microrover has a volume of 160x65x250mm (3) it weighs 3.2kg including a payload mass of 1kg and it has a peak power of 5W. The scientific model payload of the rover is a Geochemistry Instrument Package Facility (GIPF), which analyses the chemical and mineralogical composition of planetary surfaces. It consists of: An Alpha-Particle-Xray-spectrometer, a Mößbauer spectrometer and a miniature imaging system. The amount of science which can be performed within the operating range of the lander is limited since there are only few reachable, scientific interesting objects. By travelling to new sites with the aid of a microrover, the additional reach enhances the mission output and provides a significant increase in scientific return. The implementation of the Nanokhod rover on the ExoMars Lander increases its operating range by a radius of several meters, requiring only a minor mass impact of few kilograms. The Nanokhod rover is a tethered vehicle based on a Russian concept. It stays connected to the Lander via thin cables throughout the mission. This connection is used for power supply to the rover as well as the transmission of commands and scientific data. This solution minimises the communication unit and eliminates the power subsystems on the rover side, saving valuable mass and thus improving the payload to system mass ratio. By removing the power storage subsystem on the rover side, the mobile system provides a high thermal robustness and allows the system to easily survive Martian nights. The locomotion of the rover is provided by tracks. This is the optimised locomotion method on a soft and sandy surface for such a small, low-mass system, allowing even to negotiate steep slopes. The tracks enable a large contact surface of the vehicle, thus reducing its contact pressure. The sinkage is minimised reducing the bulldozing effect and increasing the traction. The central Payload Cabine has 2 (Degree of Freedom) DOF, allowing inherently robust deployment and precise payload positioning. The two drives for lifting and rotating the Payload Cabine (PLC) provides a robust and repetitive accuracy for a congruent positioning of all payload instruments on the same sample. Furthermore the PLC drives can be used for climbing and overcoming obstacles. The latest developments on the Nanokhod rover have prepared the concept for a mission scenario on the Mercury surface. The mechanical design of the Nanokhod rover was developed from a conceptual stage to an engineering model which is able to withstand the demanding conditions of the Mercury mission such as: Surface temperature of -180(°) °C, significant mass restrictions, limited power and energy supply, operational surface covered with fine dust, shock loads of 200g for 20ms and high Vacuum. With the building and testing of the engineering model the concept achieved a Technical Readiness Level (TRL) of 5 to 6, and solutions were found for a set of requirements with a high complexity. With these design requirements exceeding most mission conditions of the ExoMars lander, the current design is well-prepared for the Mars scenario.

Characterization of the room temperature payload prototype for the cryogenic interferometric gravitational wave detector KAGRA.

PubMed

Peña Arellano, Fabián Erasmo; Sekiguchi, Takanori; Fujii, Yoshinori; Takahashi, Ryutaro; Barton, Mark; Hirata, Naoatsu; Shoda, Ayaka; van Heijningen, Joris; Flaminio, Raffaele; DeSalvo, Riccardo; Okutumi, Koki; Akutsu, Tomotada; Aso, Yoichi; Ishizaki, Hideharu; Ohishi, Naoko; Yamamoto, Kazuhiro; Uchiyama, Takashi; Miyakawa, Osamu; Kamiizumi, Masahiro; Takamori, Akiteru; Majorana, Ettore; Agatsuma, Kazuhiro; Hennes, Eric; van den Brand, Jo; Bertolini, Alessandro

2016-03-01

KAGRA is a cryogenic interferometric gravitational wave detector currently under construction in the Kamioka mine in Japan. Besides the cryogenic test masses, KAGRA will also rely on room temperature optics which will hang at the bottom of vibration isolation chains. The payload of each chain comprises an optic, a system to align it, and an active feedback system to damp the resonant motion of the suspension itself. This article describes the performance of a payload prototype that was assembled and tested in vacuum at the TAMA300 site at the NAOJ in Mitaka, Tokyo. We describe the mechanical components of the payload prototype and their functionality. A description of the active components of the feedback system and their capabilities is also given. The performance of the active system is illustrated by measuring the quality factors of some of the resonances of the suspension. Finally, the alignment capabilities offered by the payload are reported.

The High-energy Burst Spectrometer for SMESE Mission

NASA Astrophysics Data System (ADS)

Gu, Qiang; Chang, Jin

2009-01-01

The SMall Explorer for Solar Eruptions (SMESE) is a small satellite being developed jointly by China and France. It is planed to launch around the next solar maximum year (˜ 2011) for observing simultaneously the two most violent types of eruptive events on the sun (the coronal mass ejection (CME) and the solar flare) and investigating their relationship. As one of the 3 main payloads of the small satellite, the high energy burst spectrometer (HEBS) adopts the upto- date high-resolution LaBr3 scintillation detector to observe the high-energy solar radiation in the range 10 keV—600 MeV. Its energy resolution is better than 3.0% at 662 keV, 2-fold higher than that of current scintillation detectors, promising a breakthrough in the studies of energy release in solar flares and CMEs, particle acceleration and the relationship between solar flares and CMEs.

Innovative Trajectory Designs to meet Exploration Challenges

NASA Technical Reports Server (NTRS)

Folta, David C.

2006-01-01

This document is a viewgraph presentation of the conference paper. Missions incorporated into NASA's Vision for Space Exploration include many different destinations and regions; are challenging to plan; and need new and innovative trajectory design methods to enable them. By combining proven methods with chaos dynamics, exploration goals that require maximum payload mass or minimum duration can be achieved. The implementation of these innovative methods, such as weak stability boundaries, has altered NASA's approach to meet exploration challenges and is described to show how exploration goals may be met in the next decade. With knowledge that various perturbations play a significant role, the mission designer must rely on both traditional design strategies as well as these innovative methods. Over the past decades, improvements have been made that would at first glance seem dramatic. This paper provides a brief narrative on how a fundamental shift has occurred and how chaos dynamics improve the design of exploration missions with complex constraints.

Power system design of ESMO

NASA Astrophysics Data System (ADS)

Ulrich, Steve; Veilleux, Jean-François; Landry Corbin, François

2009-01-01

The European Student Moon Orbiter (ESMO) spacecraft is a student-built mini satellite being designed for a mission to the Moon. Designing and launching mini satellites are becoming a current trend in the space sector since they provide an economic way to perform innovative scientific experiments and in-flight demonstration of novel space technologies. The generation, storage, control, and distribution of the electrical power in a mini satellite represents unique challenges to the power engineer since the mass and volume restrictions are very stringent. Regardless of these problems, every subsystem and payload equipment must be operated within their specified voltage band whenever they required to be turned on. This paper presents the preliminary design of a lightweight, compact, and reliable power system for ESMO that can generate 720 W. Some of the key components of the EPS include ultra triple-junction (UTJ) GaAs solar cells controlled by maximum power point trackers, and high efficiency Li-ion secondary batteries recharged in parallel.

Nonlinear robust controller design for multi-robot systems with unknown payloads

NASA Technical Reports Server (NTRS)

Song, Y. D.; Anderson, J. N.; Homaifar, A.; Lai, H. Y.

1992-01-01

This work is concerned with the control problem of a multi-robot system handling a payload with unknown mass properties. Force constraints at the grasp points are considered. Robust control schemes are proposed that cope with the model uncertainty and achieve asymptotic path tracking. To deal with the force constraints, a strategy for optimally sharing the task is suggested. This strategy basically consists of two steps. The first detects the robots that need help and the second arranges that help. It is shown that the overall system is not only robust to uncertain payload parameters, but also satisfies the force constraints.

Effects of the specular Orbiter forward radiators on a typical Spacelab payload thermal environment

NASA Technical Reports Server (NTRS)

Turner, L. D.; Humphries, W. R.; Littles, J. W.

1981-01-01

Orbiter radiators, having a specular reflection, must be considered when determining the design environment for payloads which can view the forward deployed radiators. Unlike most surfaces on the Orbiter, which reflect energy diffusely, the radiators are covered with a highly specular silverized Teflon material, with high emissivity, and have a concave contour, producing a local concentration of reflected energy towards the region of angle incidence. The combined effects of radiator specularity and geometry were analyzed using the Thermal Radiation Analysis System (TRASYS II), a specialized ray trace program, and a generalized Monte-Carlo-based thermal radiation program. Data given for a 0 deg payload inclination angle at orbital noon at 3.454 m indicate that the maximum total flux and average flux can increase 173% and 63%, respectively, when compared to diffuse radiators.

Counter-balanced, multiple cable construction crane

NASA Astrophysics Data System (ADS)

Mikulas, Martin M., Jr.; Yang, Li-Farn

1991-11-01

The invention is a counter-balanced, multiple cable construction crane. The apparatus for hoisting payloads comprises a crane having a lifting means, the lifting means comprising an end effector means and three suspension means or cables. One end of each cable attaches to a different winding means located on the lifting means, and the other end of each cable attaches to a different point on the end effector, such that the three cables have a theoretical point of convergence with this point corresponding to the center of mass of the payload. Three controls command rotation of the winding means to a predetermined position. Accordingly, the crane provides precise and autonomous positioning of the payload without human guidance. The crane further comprises a counter-balancing means. Two controls position the counter-balancing means to offset the overturning moment which arises during the lifting of heavy payloads.

Counter-balanced, multiple cable construction crane

NASA Astrophysics Data System (ADS)

Mikulas, Martin M., Jr.; Yang, Li-Farn

1993-10-01

The invention is a counter-balanced, multiple cable construction crane. The apparatus for hoisting payloads comprises a crane having a lifting means, the lifting means comprising an end effector means and three suspension means or cables. One end of each cable attaches to a different winding means located on the lifting means, and the other end of each cable attaches to a different point on the end effector, such that the three cables have a theoretical point of convergence with this point corresponding to the center of mass of the payload. Three controls command rotation of the winding means to a predetermined position. Accordingly, the crane provides precise and autonomous positioning of the payload without human guidance. The crane further comprises a counter-balancing means. Two controls position the counter-balancing means to offset the overturning moment which arises during the lifting of heavy payloads.

Counter-balanced, multiple cable construction crane

NASA Technical Reports Server (NTRS)

Mikulas, Martin M., Jr. (Inventor); Yang, Li-Farn (Inventor)

1993-01-01

The invention is a counter-balanced, multiple cable construction crane. The apparatus for hoisting payloads comprises a crane having a lifting means, the lifting means comprising an end effector means and three suspension means or cables. One end of each cable attaches to a different winding means located on the lifting means, and the other end of each cable attaches to a different point on the end effector, such that the three cables have a theoretical point of convergence with this point corresponding to the center of mass of the payload. Three controls command rotation of the winding means to a predetermined position. Accordingly, the crane provides precise and autonomous positioning of the payload without human guidance. The crane further comprises a counter-balancing means. Two controls position the counter-balancing means to offset the overturning moment which arises during the lifting of heavy payloads.

Costs and benefits of future heavy Space Freighters

NASA Astrophysics Data System (ADS)

Arend, H.

1987-10-01

A class of two-stage reusable ballistic Space Freighters with nominal launch masses of 7000 metric tons for transport of heavy payloads into low earth orbits is investigated in this paper with spcial regard to vehicle cost efficiency. A life-cycle cost analysis shows that Space Freighters with a conventional aluminum structure offer significantly lower specific transportation costs than today's systems for large payload markets and high launch rates. Advanced structural materials and thermal protection systems offer further important reductions not only with regard to vehicle mass but also with respect to specific transportation cost. A phased introduction of these technologies is cost efficient for larger programs with more than 100 vehicles.

Expendable second stage reusable space shuttle booster. Volume 4: Detail mass properties data

NASA Technical Reports Server (NTRS)

1971-01-01

Mass properties data are presented to describe the characteristics of an expendable second stage with a reusable space shuttle booster. The final mass characteristics of the vehicle configurations for three specified payloads are presented in terms of weight, center of gravity, and mass moments of inertia. Three basic subjects are the integrated vehicle system, the expendable second stage, and the booster modifications.

Experimental/analytical approaches to modeling, calibrating and optimizing shaking table dynamics for structural dynamic applications

NASA Astrophysics Data System (ADS)

Trombetti, Tomaso

This thesis presents an Experimental/Analytical approach to modeling and calibrating shaking tables for structural dynamic applications. This approach was successfully applied to the shaking table recently built in the structural laboratory of the Civil Engineering Department at Rice University. This shaking table is capable of reproducing model earthquake ground motions with a peak acceleration of 6 g's, a peak velocity of 40 inches per second, and a peak displacement of 3 inches, for a maximum payload of 1500 pounds. It has a frequency bandwidth of approximately 70 Hz and is designed to test structural specimens up to 1/5 scale. The rail/table system is mounted on a reaction mass of about 70,000 pounds consisting of three 12 ft x 12 ft x 1 ft reinforced concrete slabs, post-tensioned together and connected to the strong laboratory floor. The slip table is driven by a hydraulic actuator governed by a 407 MTS controller which employs a proportional-integral-derivative-feedforward-differential pressure algorithm to control the actuator displacement. Feedback signals are provided by two LVDT's (monitoring the slip table relative displacement and the servovalve main stage spool position) and by one differential pressure transducer (monitoring the actuator force). The dynamic actuator-foundation-specimen system is modeled and analyzed by combining linear control theory and linear structural dynamics. The analytical model developed accounts for the effects of actuator oil compressibility, oil leakage in the actuator, time delay in the response of the servovalve spool to a given electrical signal, foundation flexibility, and dynamic characteristics of multi-degree-of-freedom specimens. In order to study the actual dynamic behavior of the shaking table, the transfer function between target and actual table accelerations were identified using experimental results and spectral estimation techniques. The power spectral density of the system input and the cross power spectral density of the table input and output were estimated using the Bartlett's spectral estimation method. The experimentally-estimated table acceleration transfer functions obtained for different working conditions are correlated with their analytical counterparts. As a result of this comprehensive correlation study, a thorough understanding of the shaking table dynamics and its sensitivities to control and payload parameters is obtained. Moreover, the correlation study leads to a calibrated analytical model of the shaking table of high predictive ability. It is concluded that, in its present conditions, the Rice shaking table is able to reproduce, with a high degree of accuracy, model earthquake accelerations time histories in the frequency bandwidth from 0 to 75 Hz. Furthermore, the exhaustive analysis performed indicates that the table transfer function is not significantly affected by the presence of a large (in terms of weight) payload with a fundamental frequency up to 20 Hz. Payloads having a higher fundamental frequency do affect significantly the shaking table performance and require a modification of the table control gain setting that can be easily obtained using the predictive analytical model of the shaking table. The complete description of a structural dynamic experiment performed using the Rice shaking table facility is also reported herein. The object of this experimentation was twofold: (1) to verify the testing capability of the shaking table and, (2) to experimentally validate a simplified theory developed by the author, which predicts the maximum rotational response developed by seismic isolated building structures characterized by non-coincident centers of mass and rigidity, when subjected to strong earthquake ground motions.

Design of a ram accelerator mass launch system

NASA Technical Reports Server (NTRS)

Aarnio, Michael; Armerding, Calvin; Berschauer, Andrew; Christofferson, Erik; Clement, Paul; Gohd, Robin; Neely, Bret; Reed, David; Rodriguez, Carlos; Swanstrom, Fredrick

1988-01-01

The ram accelerator mass launch system has been proposed to greatly reduce the costs of placing acceleration-insensitive payloads into low earth orbit. The ram accelerator is a chemically propelled, impulsive mass launch system capable of efficiently accelerating relatively large masses from velocities of 0.7 km/sec to 10 km/sec. The principles of propulsion are based on those of a conventional supersonic air-breathing ramjet; however the device operates in a somewhat different manner. The payload carrying vehicle resembles the center-body of the ramjet and accelerates through a stationary tube which acts as the outer cowling. The tube is filled with premixed gaseous fuel and oxidizer mixtures that burn in the vicinity of the vehicle's base, producing a thrust which accelerates the vehicle through the tube. This study examines the requirement for placing a 2000 kg vehicle into a 500 km circular orbit with a minimum amount of on-board rocket propellant for orbital maneuvers. The goal is to achieve a 50 pct payload mass fraction. The proposed design requirements have several self-imposed constraints that define the vehicle and tube configurations. Structural considerations on the vehicle and tube wall dictate an upper acceleration limit of 1000 g's and a tube inside diameter of 1.0 m. In-tube propulsive requirements and vehicle structural constraints result in a vehicle diameter of 0.76 m, a total length of 7.5 m and a nose-cone half angle of 7 degrees. An ablating nose-cone constructed from carbon-carbon composite serves as the thermal protection mechanism for atmospheric transit.

A Versatile Lifting Device for Lunar Surface Payload Handling, Inspection and Regolith Transport Operations

NASA Technical Reports Server (NTRS)

Doggett, William R.; Dorsey, John T.; Collins, Timothy J.; King, Bruce D.; Mikulas, Martin M., Jr.

2008-01-01

Devices for lifting and transporting payloads and material are critical for efficient Earth-based construction operations. Devices with similar functionality will be needed to support lunar-outpost construction, servicing, inspection, regolith excavation, grading and payload placement. Past studies have proposed that only a few carefully selected devices are required for a lunar outpost. One particular set of operations involves lifting and manipulating payloads in the 100 kg to 3,000 kg range, which are too large or massive to be handled by unassisted astronauts. This paper will review historical devices used for payload handling in space and on earth to derive a set of desirable features for a device that can be used on planetary surfaces. Next, an innovative concept for a lifting device is introduced, which includes many of the desirable features. The versatility of the device is discussed, including its application to lander unloading, servicing, inspection, regolith excavation and site preparation. Approximate rules, which can be used to size the device for specific payload mass and reach requirements, are provided. Finally, details of a test-bed implementation of the innovative concept, which will be used to validate the structural design and develop operational procedures, is provided.

Simulating Vibrations in a Complex Loaded Structure

NASA Technical Reports Server (NTRS)

Cao, Tim T.

2005-01-01

The Dynamic Response Computation (DIRECT) computer program simulates vibrations induced in a complex structure by applied dynamic loads. Developed to enable rapid analysis of launch- and landing- induced vibrations and stresses in a space shuttle, DIRECT also can be used to analyze dynamic responses of other structures - for example, the response of a building to an earthquake, or the response of an oil-drilling platform and attached tanks to large ocean waves. For a space-shuttle simulation, the required input to DIRECT includes mathematical models of the space shuttle and its payloads, and a set of forcing functions that simulates launch and landing loads. DIRECT can accommodate multiple levels of payload attachment and substructure as well as nonlinear dynamic responses of structural interfaces. DIRECT combines the shuttle and payload models into a single structural model, to which the forcing functions are then applied. The resulting equations of motion are reduced to an optimum set and decoupled into a unique format for simulating dynamics. During the simulation, maximum vibrations, loads, and stresses are monitored and recorded for subsequent analysis to identify structural deficiencies in the shuttle and/or payloads.

KSC-00pp1734

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, moves higher toward the payload changeout room (PCR). The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure (RSS) (at left) that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and solid rocket boosters showing behind it. When the RSS is closed around Endeavour, the P6 truss will be able to be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC00pp1734

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, moves higher toward the payload changeout room (PCR). The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure (RSS) (at left) that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and solid rocket boosters showing behind it. When the RSS is closed around Endeavour, the P6 truss will be able to be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC00pp1735

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, is lifted higher toward the payload changeout room (PCR) above it. The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure (RSS) (at left) that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and one solid rocket booster showing behind it. When the RSS is closed around Endeavour, the P6 truss will be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC-00pp1733

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, is lifted toward the payload changeout room (PCR). The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure (RSS) (on the left) that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and one solid rocket booster showing behind it. When the RSS is closed around Endeavour, the P6 truss will be able to be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC00pp1733

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, is lifted toward the payload changeout room (PCR). The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure (RSS) (on the left) that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and one solid rocket booster showing behind it. When the RSS is closed around Endeavour, the P6 truss will be able to be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

KSC-00pp1735

NASA Image and Video Library

2000-11-14

KENNEDY SPACE CENTER, FLA. -- On Launch Pad 39B, the payload transport canister, with the P6 integrated truss segment inside, is lifted higher toward the payload changeout room (PCR) above it. The PCR is the enclosed, environmentally controlled portion of the Rotating Service Structure (RSS) (at left) that supports payload delivery at the pad. At right is Space Shuttle Endeavour with its orange external tank and one solid rocket booster showing behind it. When the RSS is closed around Endeavour, the P6 truss will be moved into the orbiter’s payload bay. The P6, payload on mission STS-97, comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

Design of a Ram Accelerator mass launch system

NASA Technical Reports Server (NTRS)

1988-01-01

The Ram Accelerator, a chemically propelled, impulsive mass launch system, is presented as a viable concept for directly launching acceleration-insensitive payloads into low Earth orbit. The principles of propulsion are based on those of an airbreathing supersonic ramjet. The payload vehicle acts as the ramjet centerbody and travels through a fixed launch tube that acts as the ramjet outer cowling. The launch tube is filled with premixed gaseous fuel and oxidizer mixtures that combust at the base of the vehicle and produce thrust. Two modes of in-tube propulsion involving ramjet cycles are used in sequence to accelerate the vehicle from 0.7 km/sec to 9 km/sec. Requirements for placing a 2000 kg vehicle into a 500-km circular orbit, with a minimum amount of onboard rocket propellant for orbital maneuvers, are examined. It is shown that in-tube propulsion requirements dictate a launch tube length of 5.1 km to achieve an exit velocity of 9 km/sec, with peak accelerations not to exceed 1000 g's. Aerodynamic heating due to atmospheric transit requires minimal ablative protection and the vehicle retains a large percentage of its exit velocity. An indirect orbital insertion maneuver with aerobraking and two apogee burns is examined to minimize the required onboard propellant mass. An appropriate onboard propulsion system design to perform the required orbital maneuvers with minimum mass requirements is also determined. The structural designs of both the launch tube and the payload vehicle are examined using simple structural and finite element analysis for various materials.

Inflation of Unreefed and Reefed Extraction Parachutes

NASA Technical Reports Server (NTRS)

Ray, Eric S.; Varela, Jose G.

2015-01-01

Data from the Orion and several other test programs have been used to reconstruct inflation parameters for 28 ft Do extraction parachutes as well as the parent aircraft pitch response during extraction. The inflation force generated by extraction parachutes is recorded directly during tow tests but is usually inferred from the payload accelerometer during Low Velocity Airdrop Delivery (LVAD) flight test extractions. Inflation parameters are dependent on the type of parent aircraft, number of canopies, and standard vs. high altitude extraction conditions. For standard altitudes, single canopy inflations are modeled as infinite mass, but the non-symmetric inflations in a cluster are modeled as finite mass. High altitude extractions have necessitated reefing the extraction parachutes, which are best modeled as infinite mass for those conditions. Distributions of aircraft pitch profiles and inflation parameters have been generated for use in Monte Carlo simulations of payload extractions.

A procedure obtaining stiffnesses and masses of a structure from vibration modes and substructure static test data

NASA Technical Reports Server (NTRS)

Edighoffer, H. H.

1979-01-01

A component mode desynthesis procedure is developed for determining the unknown vibration characteristics of a structural component (i.e., a launch vehicle) given the vibration characteristics of a structural system composed of that component combined with a known one (i.e., a payload). At least one component static test has to be performed. These data are used in conjunction with the system measured frequencies and mode shapes to obtain the vibration characteristics of each component. The flight dynamics of an empty launch vehicle can be determined from measurements made on a vehicle/payload combination in conjunction with a static test on the payload.

Advanced APS impacts on vehicle payloads

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Reed, Brian D.

1989-01-01

Advanced auxiliary propulsion system (APS) technology has the potential to both, increase the payload capability of earth-to-orbit (ETO) vehicles by reducing APS propellant mass, and simplify ground operations and logistics by reducing the number of fluids on the vehicle and eliminating toxic, corrosive propellants. The impact of integrated cryogenic APS on vehicle payloads is addressed. In this system, launch propulsion system residuals are scavenged from integral launch propulsion tanks for use in the APS. Sufficient propellant is preloaded into the APS to return to earth with margin and noncomplete scavenging assumed. No propellant conditioning is required by the APS, but ambient heat soak is accommodated. High temperature rocket materials enable the use of the unconditioned hydrogen/oxygen in the APS and are estimated to give APS rockets specific impulse of up to about 444 sec. The payload benefits are quantified and compared with an uprated monomethylhydrazine/nitrogen tetroxide system in a conservative fashion, by assuming a 25.5 percent weight growth for the hydrogen/oxygen system and a 0 percent weight growth for the uprated system. The combination of scavenging and high performance gives payload impacts which are highly mission specific. A payload benefit of 861 kg (1898 lbm) was estimated for a Space Station Freedom rendezvous mission and 2099 kg (4626 lbm) for a sortie mission, with payload impacts varying with the amount of launch propulsion residual propellants. Missions without liquid propellant scavenging were estimated to have payload penalties, however, operational benefits were still possible.

Advanced APS Impacts on Vehicle Payloads

NASA Technical Reports Server (NTRS)

Schneider, Steven J.; Reed, Brian D.

1989-01-01

Advanced auxiliary propulsion system (APS) technology has the potential to both, increase the payload capability of earth-to-orbit (ETO) vehicles by reducing APS propellant mass, and simplify ground operations and logistics by reducing the number of fluids on the vehicle and eliminating toxic, corrosive propellants. The impact of integrated cryogenic APS on vehicle payloads is addressed. In this system, launch propulsion system residuals are scavenged from integral launch propulsion tanks for use in the APS. Sufficient propellant is preloaded into the APS to return to earth with margin and noncomplete scavenging assumed. No propellant conditioning is required by the APS, but ambient heat soak is accommodated. High temperature rocket materials enable the use of the unconditioned hydrogen/oxygen in the APS and are estimated to give APS rockets specific impulse of up to about 444 sec. The payload benefits are quantified and compared with an uprated monomethyl hydrazine/nitrogen tetroxide system in a conservative fashion, by assuming a 25.5 percent weight growth for the hydrogen/oxygen system and a 0 percent weight growth for the uprated system. The combination and scavenging and high performance gives payload impacts which are highly mission specific. A payload benefit of 861 kg (1898 lbm) was estimated for a Space Station Freedom rendezvous mission and 2099 kg (4626 lbm) for a sortie mission, with payload impacts varying with the amount of launch propulsion residual propellants. Missions without liquid propellant scavenging were estimated to have payload penalties, however, operational benefits were still possible.

Column Number Density Expressions Through M = 0 and M = 1 Point Source Plumes Along Any Straight Path

NASA Technical Reports Server (NTRS)

Woronowicz, Michael

2017-01-01

Providers of payloads carried aboard the International Space Station must conduct analyses to demonstrate that any planned gaseous venting events generate no more than a certain level of material that may interfere with optical measurements from other experiments or payloads located nearby. This requirement is expressed in terms of a maximum column number density (CND). Depending on the level of rarefaction, such venting may be characterized by effusion for low flow rates, or by a sonic distribution at higher levels. Since the relative locations of other sensitive payloads are often unknown because they may refer to future projects, this requirement becomes a search for the maximum CND along any path.In another application, certain astronomical observations make use of CND to estimate light attenuation from a distant star through gaseous plumes, such as the Fermi Bubbles emanating from the vicinity of the black hole at the center of our Milky Way galaxy, in order to infer the amount of material being expelled via those plumes.This paper presents analytical CND expressions developed for general straight paths based upon a free molecule point source model for steady effusive flow and for a distribution fitted to model flows from a sonic orifice. Among other things, in this Mach number range it is demonstrated that the maximum CND from a distant location occurs along the path parallel to the source plane that intersects the plume axis. For effusive flows this value is exactly twice the CND found along the ray originating from that point of intersection and extending to infinity along the plumes axis. For sonic plumes this ratio is reduced to about 43.

Science and Applications Space Platform (SASP) End-to-End Data System Study

NASA Technical Reports Server (NTRS)

Crawford, P. R.; Kasulka, L. H.

1981-01-01

The capability of present technology and the Tracking and Data Relay Satellite System (TDRSS) to accommodate Science and Applications Space Platforms (SASP) payload user's requirements, maximum service to the user through optimization of the SASP Onboard Command and Data Management System, and the ability and availability of new technology to accommodate the evolution of SASP payloads were assessed. Key technology items identified to accommodate payloads on a SASP were onboard storage devices, multiplexers, and onboard data processors. The primary driver is the limited access to TDRSS for single access channels due to sharing with all the low Earth orbit spacecraft plus shuttle. Advantages of onboard data processing include long term storage of processed data until TRDSS is accessible, thus reducing the loss of data, eliminating large data processing tasks at the ground stations, and providing a more timely access to the data.

Performance analysis of a laser propelled interorbital tansfer vehicle

NASA Technical Reports Server (NTRS)

Minovitch, M. A.

1976-01-01

Performance capabilities of a laser-propelled interorbital transfer vehicle receiving propulsive power from one ground-based transmitter was investigated. The laser transmits propulsive energy to the vehicle during successive station fly-overs. By applying a series of these propulsive maneuvers, large payloads can be economically transferred between low earth orbits and synchronous orbits. Operations involving the injection of large payloads onto escape trajectories are also studied. The duration of each successive engine burn must be carefully timed so that the vehicle reappears over the laser station to receive additional propulsive power within the shortest possible time. The analytical solution for determining these time intervals is presented, as is a solution to the problem of determining maximum injection payloads. Parameteric computer analysis based on these optimization studies is presented. The results show that relatively low beam powers, on the order of 50 MW to 60 MW, produce significant performance capabilities.

Innovative mechanism for measuring the mass properties of an object

NASA Technical Reports Server (NTRS)

Wolcott, Kedron R.; Graham, Todd A.; Doty, Keith L.

1994-01-01

The Kennedy Space Center Robotics Group recently completed development and testing on a novel approach to measure the mass properties of a rigid body. This unique design can measure the payload's weight, mass center location, and moments of inertia about three orthogonal axes. Furthermore, these measurements only require a single torque sensor and a single angular position sensor.

Antibody Drug Conjugates: Application of Quantitative Pharmacology in Modality Design and Target Selection.

PubMed

Sadekar, S; Figueroa, I; Tabrizi, M

2015-07-01

Antibody drug conjugates (ADCs) are a multi-component modality comprising of an antibody targeting a cell-specific antigen, a potent drug/payload, and a linker that can be processed within cellular compartments to release payload upon internalization. Numerous ADCs are being evaluated in both research and clinical settings within the academic and pharmaceutical industry due to their ability to selectively deliver potent payloads. Hence, there is a clear need to incorporate quantitative approaches during early stages of drug development for effective modality design and target selection. In this review, we describe a quantitative approach and framework for evaluation of the interplay between drug- and systems-dependent properties (i.e., target expression, density, localization, turnover, and affinity) in order to deliver a sufficient amount of a potent payload into the relevant target cells. As discussed, theoretical approaches with particular considerations given to various key properties for the target and modality suggest that delivery of the payload into particular effect cells to be more sensitive to antigen concentrations for targets with slow turnover rates as compared to those with faster internalization rates. Further assessments also suggest that increasing doses beyond the threshold of the target capacity (a function of target internalization and expression) may not impact the maximum amount of payload delivered to the intended effect cells. This article will explore the important application of quantitative sciences in selection of the target and design of ADC modalities.

The Cost-Optimal Size of Future Reusable Launch Vehicles

NASA Astrophysics Data System (ADS)

Koelle, D. E.

2000-07-01

The paper answers the question, what is the optimum vehicle size — in terms of LEO payload capability — for a future reusable launch vehicle ? It is shown that there exists an optimum vehicle size that results in minimum specific transportation cost. The optimum vehicle size depends on the total annual cargo mass (LEO equivalent) enviseaged, which defines at the same time the optimum number of launches per year (LpA). Based on the TRANSCOST-Model algorithms a wide range of vehicle sizes — from 20 to 100 Mg payload in LEO, as well as launch rates — from 2 to 100 per year — have been investigated. It is shown in a design chart how much the vehicle size as well as the launch rate are influencing the specific transportation cost (in MYr/Mg and USS/kg). The comparison with actual ELVs (Expendable Launch Vehicles) and Semi-Reusable Vehicles (a combination of a reusable first stage with an expendable second stage) shows that there exists only one economic solution for an essential reduction of space transportation cost: the Fully Reusable Vehicle Concept, with rocket propulsion and vertical take-off. The Single-stage Configuration (SSTO) has the best economic potential; its feasibility is not only a matter of technology level but also of the vehicle size as such. Increasing the vehicle size (launch mass) reduces the technology requirements because the law of scale provides a better mass fraction and payload fraction — practically at no cost. The optimum vehicle design (after specification of the payload capability) requires a trade-off between lightweight (and more expensive) technology vs. more conventional (and cheaper) technology. It is shown that the the use of more conventional technology and accepting a somewhat larger vehicle is the more cost-effective and less risky approach.

Ultra-light weight undamped tuned dynamic absorber for cryogenically cooled infrared electro-optic payload

NASA Astrophysics Data System (ADS)

Veprik, Alexander; Babitsky, Vladimir

2017-04-01

Attenuation of tonal cryocooler induced vibration in infrared electro-optical payloads may be achieved by using of Tuned Dynamic Absorber (TDA) which is, generally speaking, a passive, weakly damped mass-spring system the resonant frequency of which is precisely matched with the driving frequency. Added TDA results in a favorable modification of the frequency response functions of combined structure. In particular, a favorable antiresonant notch appears at the frequency of tonal excitation along with the adjacent secondary resonance, the width and depth of which along with its closeness to the secondary resonance are strongly dependent on the mass and damping ratios. Using heavier TDA favorably results in wider and deeper antiresonant notch along with increased gap between antiresonant and resonant frequencies. Lowering damping in TDA favorably results in deepening the antiresonant notch. The weight of TDA is usually subjected to tight design constrains. Use of lightweight TDA not only diminishes the attainable performance but also complicates the procedure of frequency matching. Along these lines, even minor frequency deviations may negate the TDA performance and even result in TDA failure in case of resonant build up. The authors are presenting theoretical and practical aspects of designing and constructing ultra-light weight TDA in application to vibration attenuation of electro-optical infrared payload relying on Split Stirling linear cryocooler, the driving frequency of which is fixed and may be accurately tuned and maintained using a digital controller over the entire range of working conditions and lifetime; the lack of mass ratio is compensated by minimizing the damping ratio. In one particular case, in excess of 100-fold vibration attenuation has been achieved by adding as little as 5% to the payload weight.

Low-Cost Propellant Launch From a Tethered Balloon

NASA Technical Reports Server (NTRS)

Wilcox, Brian

2006-01-01

A document presents a concept for relatively inexpensive delivery of propellant to a large fuel depot in low orbit around the Earth, for use in rockets destined for higher orbits, the Moon, and for remote planets. The propellant is expected to be at least 85 percent of the mass needed in low Earth orbit to support the NASA Exploration Vision. The concept calls for the use of many small ( 10 ton) spin-stabilized, multistage, solid-fuel rockets to each deliver 250 kg of propellant. Each rocket would be winched up to a balloon tethered above most of the atmospheric mass (optimal altitude 26 2 km). There, the rocket would be aimed slightly above the horizon, spun, dropped, and fired at a time chosen so that the rocket would arrive in orbit near the depot. Small thrusters on the payload (powered, for example, by boil-off gases from cryogenic propellants that make up the payload) would precess the spinning rocket, using data from a low-cost inertial sensor to correct for small aerodynamic and solid rocket nozzle misalignment torques on the spinning rocket; would manage the angle of attack and the final orbit insertion burn; and would be fired on command from the depot in response to observations of the trajectory of the payload so as to make small corrections to bring the payload into a rendezvous orbit and despin it for capture by the depot. The system is low-cost because the small rockets can be mass-produced using the same techniques as those to produce automobiles and low-cost munitions, and one or more can be launched from a U.S. territory on the equator (Baker or Jarvis Islands in the mid-Pacific) to the fuel depot on each orbit (every 90 minutes, e.g., any multiple of 6,000 per year).

An air-breathing ballistic space transporter for Europe

NASA Technical Reports Server (NTRS)

Kramer, P. A.; Buehler, R. D.

1985-01-01

With increasing transport requirements, reusable space transporters again receive serious consideration in Europe as successors to the Ariane family. The paper deals with a hydrogen-ramjet-propelled, 1-1/2-stage reusable ballistic space transporter with vertical take-off and landing and using liquid hydrogen/oxygen rockets. This novel concept was developed in a theoretical study at the University of Stuttgart. The results are compared with recently published studies of several other European space transporter concepts. The data derived for the Istra - concept are: 15.4 Mg payload into low Earth-orbit, 155 Mg gross lift-off mass, 10% payload ratio, which represents a 57% propellant saving, and 44% reduction in dry mass (structure and engines) compared with comparable two-stage pure rocket concepts.

Tether System for Exchanging Payloads Between the International Space Station and the Lunar Surface

NASA Technical Reports Server (NTRS)

Hoyt, Robert P.

1998-01-01

Systems composed of several rotating and/or hanging tethers may provide a means of exchanging supplies between low Earth orbit facilities and lunar bases without requiring the use of propellant. This work develops methods for designing a tether system capable of repeatedly exchanging payloads between a LEO facility such as the International Space Station or a Space Business Park and a base on the lunar surface. In this system, a hanging tether extended upwards from the LEO facility, places a payload into a slightly elliptical orbit, where it is caught by a rotating tether in a higher elliptical orbit. This rotating tether then tosses the payload to the moon. At the moon, a long rotating "Lunavator" tether catches the payload and deposits it on the surface of the moon. By transporting an equal mass of lunar materials such as oxygen back down to the LEO facility through the tether transport system, the momentum and energy of the system is conserved, allowing frequent traffic between LEO and the lunar surface with minimal propellant requirements.

Space science/space station attached payload pointing accommodation study: Technology assessment white paper

NASA Technical Reports Server (NTRS)

Lin, Richard Y.; Mann, Kenneth E.; Laskin, Robert A.; Sirlin, Samuel W.

1987-01-01

Technology assessment is performed for pointing systems that accommodate payloads of large mass and large dimensions. Related technology areas are also examined. These related areas include active thermal lines or power cables across gimbals, new materials for increased passive damping, tethered pointing, and inertially reacting pointing systems. Conclusions, issues and concerns, and recommendations regarding the status and development of large pointing systems for space applications are made based on the performed assessments.

Dobson space telescope: development of an optical payload of the next generation

NASA Astrophysics Data System (ADS)

Segert, Tom; Danziger, Björn; Gork, Daniel; Lieder, Matthias

2017-11-01

The Dobson Space Telescope (DST) is a research project of the Department of Astronautics at the TUBerlin. For Development and commercialisation there is a close cooperation with the network of the Berlin Space Industry (RIBB). Major Partner is the Astro- und Feinwerktechnik Adlershof GmbH a specialist for space structures and head of the industry consortia which built the DLR BIRD micro satellite. The aim of the project is to develop a new type of deployable telescope that can overcome the mass and volume limitations of small satellites. With the DST payload micro satellites of the 100kg class will be able to carry 50cm main mirror diameter optics (→ 1m GSD). Basis of this technology is the fact that a telescope is mainly empty space between the optical elements. To fold down the telescope during launch and to undfold it after the satellite reached its orbit can save 70% of payload volume and 50% of payload mass. Since these advantages continue along the value added chain DST is of highest priority for the next generation of commercial EO micro satellites. Since 2002 the key technologies for DST have been developed in test benches in Labs of TU-Berlin and were tested on board a ESA parabolic flight campaign in 2005. The development team at TU-Berlin currently prepares the foundation of a start-up company for further development and commercialisation of DST.

Multi-spectral optical scanners for commercial earth observation missions

NASA Astrophysics Data System (ADS)

Schröter, Karin; Engel, Wolfgang; Berndt, Klaus

2017-11-01

In recent years, a number of commercial Earth observation missions have been initiated with the aim to gather data in the visible and near-infrared wavelength range. Some of these missions aim at medium resolution (5 to 10 m) multi-spectral imaging with the special background of daily revisiting. Typical applications aim at monitoring of farming area for growth control and harvest prediction, irrigation control, or disaster monitoring such as hail damage in farming, or flood survey. In order to arrive at profitable business plans for such missions, it is mandatory to establish the space segment, i.e. the spacecraft with their opto -electronic payloads, at minimum cost while guaranteeing maximum reliability for mission success. As multiple spacecraft are required for daily revisiting, the solutions are typically based on micro-satellites. This paper presents designs for multi-spectral opto-electric scanners for this type of missions. These designs are drive n by minimum mass and power budgets of microsatellites, and the need for minimum cost. As a consequence, it is mandatory to arrive at thermally robust, compact telescope designs. The paper gives a comparison between refractive, catadioptric, and TMA optics. For mirror designs, aluminium and Zerodur mirror technologies are briefly discussed. State-of-the art focal plane designs are presented. The paper also addresses the choice of detector technologies such as CCDs and CMOS Active Pixel Sensors. The electronics of the multi-spectral scanners represent the main design driver regarding power consumption, reliability, and (most often) cost. It can be subdivided into the detector drive electronics, analog and digital data processing chains, the data mass memory unit, formatting and down - linking units, payload control electronics, and local power supply. The paper gives overviews and trade-offs between data compression strategies and electronics solutions, mass memory unit designs, and data formatting approaches. Special emphasis will be put on space application aspects of these electronics solutions such as radiation total dose tolerance and single events robustness. Finally, software architecture and operational modes of commercial multi-spectral scanners are discussed. They are driven by operational requirements and mission constraints such as data takes per orbit, number of downlink ground stations, calibration needs, and mission schedule planning.

Boeing-747 aircraft with external cargo pod

NASA Technical Reports Server (NTRS)

Quartero, C. B.; Washburn, G. F.; Price, J. E.

1978-01-01

An analysis was conducted to investigate the feasibility of mounting a detachable pod to the underside of the fuselage of a Boeing Model 747 aircraft to carry outsized cargo in case of military emergency. The analysis showed that the 747 configured with the pod and carrying only a bridge launcher as payload attained a range of 8.70 Mm (4 700 n. mi.) at Mach .68. This range was based on a maximum take-off gross weight of 3.447 MN (775 000 1bf) which included 212 kN (47 700 lbf) pod weight and 543 kN (122 000 lbf) payload (bridge launcher).

Optimal three-dimensional reusable tug trajectories for planetary missions including correction for nodal precession

NASA Technical Reports Server (NTRS)

Borsody, J.

1976-01-01

Equations are derived by using the maximum principle to maximize the payload of a reusable tug for planetary missions. The analysis includes a correction for precession of the space shuttle orbit. The tug returns to this precessed orbit (within a specified time) and makes the required nodal correction. A sample case is analyzed that represents an inner planet mission as specified by a fixed declination and right ascension of the outgoing asymptote and the mission energy. The reusable stage performance corresponds to that of a typical cryogenic tug. Effects of space shuttle orbital inclination, several trajectory parameters, and tug thrust on payload are also investigated.

A contribution to the availability of lunar resources for powered construction

NASA Technical Reports Server (NTRS)

Heppenheimer, T. A.

1980-01-01

The use of lunar resources to construct solar power satellites wherein the resources are transported by a lunar mass driver is discussed. The minimization of cross track errors in the launch of payloads by mass driver is emphasized. The design and construction of the mass driver is outlined. Features of the proposed system addressed include passive magnetic damping, separation and snapout, and downrange correction.

An evaluation of composite propulsion for single-stage-to-orbit vehicles designed for horizontal take-off

NASA Technical Reports Server (NTRS)

Martin, J. A.

1977-01-01

Composite propulsion was analyzed for single-stage-to-orbit vehicles designed for horizontal take-off. Trajectory, geometric, and mass analyses were performed to establish the orbital payload capability of six engines. The results indicated that none of the engines performed adequately to deliver payloads to orbit as analyzed. The single-stage turbine and oxidizer-rich gas generator resulted in a low engine specific impulse, and the performance increment of the ejector subsystem was less than that of a separate rocket system with a high combustion pressure. There was a benefit from incorporating a fan into the engine, and removal of the fan from the airstream during the ramjet mode increased the orbital payload capability.

Area V: A National Launch Asset for the 21st Century

NASA Technical Reports Server (NTRS)

Sumrall, Phil

2009-01-01

The goal of this presentation is to present an update on status and development of the Ares V launch vehicle. The Ares V is a heavy lift vehicle that is being designed to launch cargo into Low Earth Orbit and transfer Cargo and crews to the Moon. Slides show the commonalities between the Ares V, and the Ares I, and the Delta IV. The launch profile for a typical Lunar mission is reviewed. A timeline showing the progress from the Exploration Systems Architecture Study (ESAS) to the Lunar Capability Concept Review (LCCR) is presented. Other slides review the payload shroud, the payload vs altitude and inclination, the payload mass vs C3 Energy, projections of the performance for selected trajectories, and the planning calendar.

Adaptive tracking for complex systems using reduced-order models

NASA Technical Reports Server (NTRS)

Carnigan, Craig R.

1990-01-01

Reduced-order models are considered in the context of parameter adaptive controllers for tracking workspace trajectories. A dual-arm manipulation task is used to illustrate the methodology and provide simulation results. A parameter adaptive controller is designed to track a payload trajectory using a four-parameter model instead of the full-order, nine-parameter model. Several simulations with different payload-to-arm mass ratios are used to illustrate the capabilities of the reduced-order model in tracking the desired trajectory.

Adaptive tracking for complex systems using reduced-order models

NASA Technical Reports Server (NTRS)

Carignan, Craig R.

1990-01-01

Reduced-order models are considered in the context of parameter adaptive controllers for tracking workspace trajectories. A dual-arm manipulation task is used to illustrate the methodology and provide simulation results. A parameter adaptive controller is designed to track the desired position trajectory of a payload using a four-parameter model instead of a full-order, nine-parameter model. Several simulations with different payload-to-arm mass ratios are used to illustrate the capabilities of the reduced-order model in tracking the desired trajectory.

SAM-CAAM: A Concept for Acquiring Systematic Aircraft Measurements to Characterize Aerosol Air Masses.

PubMed

Kahn, Ralph A; Berkoff, Tim A; Brock, Charles; Chen, Gao; Ferrare, Richard A; Ghan, Steven; Hansico, Thomas F; Hegg, Dean A; Martins, J Vanderlei; McNaughton, Cameron S; Murphy, Daniel M; Ogren, John A; Penner, Joyce E; Pilewskie, Peter; Seinfeld, John H; Worsnop, Douglas R

2017-10-01

A modest operational program of systematic aircraft measurements can resolve key satellite-aerosol-data-record limitations. Satellite observations provide frequent, global aerosol-amount maps, but offer only loose aerosol property constraints needed for climate and air quality applications. We define and illustrate the feasibility of flying an aircraft payload to measure key aerosol optical, microphysical, and chemical properties in situ . The flight program could characterize major aerosol air-mass types statistically, at a level-of-detail unobtainable from space. It would: (1) enhance satellite aerosol retrieval products with better climatology assumptions, and (2) improve translation between satellite-retrieved optical properties and species-specific aerosol mass and size simulated in climate models to assess aerosol forcing, its anthropogenic components, and other environmental impacts. As such, Systematic Aircraft Measurements to Characterize Aerosol Air Masses (SAM-CAAM) could add value to data records representing several decades of aerosol observations from space, improve aerosol constraints on climate modeling , help interrelate remote-sensing, in situ, and modeling aerosol-type definitions , and contribute to future satellite aerosol missions. Fifteen Required Variables are identified, and four Payload Options of increasing ambition are defined, to constrain these quantities. "Option C" could meet all the SAM-CAAM objectives with about 20 instruments, most of which have flown before, but never routinely several times per week, and never as a group. Aircraft integration, and approaches to data handling, payload support, and logistical considerations for a long-term, operational mission are discussed. SAM-CAAM is feasible because, for most aerosol sources and specified seasons, particle properties tend to be repeatable , even if aerosol loading varies.

A preliminary investigation of the environmental Control and Life Support Subsystems (EC/LSS) for animal and plant experiment payloads

NASA Technical Reports Server (NTRS)

Wells, H. B.

1972-01-01

A preliminary study of the environmental control and life support subsystems (EC/LSS) necessary for an earth orbital spacecraft to conduct biological experiments is presented. The primary spacecraft models available for conducting these biological experiments are the space shuttle and modular space station. The experiments would be housed in a separate module that would be contained in either the shuttle payload bay or attached to the modular space station. This module would be manned only for experiment-related tasks, and would contain a separate EC/LSS for the crew and animals. Metabolic data were tabulated on various animals that are considered useful for a typical experiment program. The minimum payload for the 30-day space shuttle module was found to require about the equivalent of a one-man EC/LSS; however, the selected two-man shuttle assemblies will give a growth and contingency factor of about 50 percent. The maximum payloads for the space station mission will require at least a seven-man EC/LSS for the laboratory colony and a nine-man EC/LSS for the centrifuge colony. There is practically no room for growth or contingencies in these areas.

14 CFR 125.271 - Emergency and emergency evacuation duties.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Emergency and emergency evacuation duties. 125.271 Section 125.271 Aeronautics and Space FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF... CERTIFICATION AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD...

14 CFR 125.73 - Contents.

Code of Federal Regulations, 2010 CFR

2006-01-01

... HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE... include— (a) The name of each management person who is authorized to act for the certificate holder, the...) Procedures for ensuring compliance with airplane weight and balance limitations; (c) Copies of the...

14 CFR 125.407 - Maintenance log: Airplanes.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Maintenance log: Airplanes. 125.407 Section... OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6... Maintenance log: Airplanes. (a) Each person who takes corrective action or defers action concerning a reported...

14 CFR 125.407 - Maintenance log: Airplanes.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Maintenance log: Airplanes. 125.407 Section... OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6... Maintenance log: Airplanes. (a) Each person who takes corrective action or defers action concerning a reported...

14 CFR 125.91 - Airplane requirements: General.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Airplane requirements: General. 125.91... AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements...

14 CFR 125.91 - Airplane requirements: General.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Airplane requirements: General. 125.91... AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements...

14 CFR 125.407 - Maintenance log: Airplanes.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Maintenance log: Airplanes. 125.407 Section... OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6... Maintenance log: Airplanes. (a) Each person who takes corrective action or defers action concerning a reported...

14 CFR 125.91 - Airplane requirements: General.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Airplane requirements: General. 125.91... AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements...

14 CFR 125.407 - Maintenance log: Airplanes.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Maintenance log: Airplanes. 125.407 Section... OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6... Maintenance log: Airplanes. (a) Each person who takes corrective action or defers action concerning a reported...

14 CFR 125.91 - Airplane requirements: General.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Airplane requirements: General. 125.91... AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements...

14 CFR 125.91 - Airplane requirements: General.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane requirements: General. 125.91... AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements...

14 CFR 125.407 - Maintenance log: Airplanes.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Maintenance log: Airplanes. 125.407 Section... OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6... Maintenance log: Airplanes. (a) Each person who takes corrective action or defers action concerning a reported...

Optical engine initiation: multiple compartment applications

NASA Astrophysics Data System (ADS)

Hunt, Jeffrey H.

2009-05-01

Modern day propulsion systems are used in aerospace applications for different purposes. The aerospace industry typically requires propulsion systems to operate in a rocket mode in order to drive large boost vehicles. The defense industry generally requires propulsion systems to operate in an air-breathing mode in order to drive missiles. A mixed system could use an air-breathing first stage and a rocket-mode upper stage for space access. Thus, propulsion systems can be used for high mass payloads and where the payload is dominated by the fuel/oxidizer mass being used by the propulsion system. The pulse detonation wave engine (PDWE) uses an alternative type of detonation cycle to achieve the same propulsion results. The primary component of the PDWE is the combustion chamber (or detonation tube). The PDWE represents an attractive propulsion source since its engine cycle is thermodynamically closest to that of a constant volume reaction. This characteristic leads to the inference that a maximum of the potential energy of the PDWE is put into thrust and not into flow work. Consequently, the volume must be increased. The technical community has increasingly adopted the alternative choice of increasing total volume by designing the engine to include a set of banks of smaller combustion chambers. This technique increases the complexity of the ignition subsystem because the inter-chamber timing must be considered. Current approaches to igniting the PDWE have involved separate shock or blast wave initiators and chemical additives designed to enhance detonatibility. An optical ignition subsystem generates a series of optical pulses, where the optical pulses ignite the fuel/oxidizer mixture such that the chambers detonate in a desired order. The detonation system also has an optical transport subsystem for transporting the optical pulses from the optical ignition subsystem to the chambers. The use of optical ignition and transport provides a non-toxic, small, lightweight, precisely controlled detonation system.

NASA/USRA advanced space design program: The laser powered interorbital vehicle

NASA Technical Reports Server (NTRS)

1989-01-01

A preliminary design is presented for a low-thrust Laser Powered Interorbital Vehicle (LPIV) intended for cargo transportation between an earth space station and a lunar base. The LPIV receives its power from two iodide laser stations, one orbiting the earth and the other located on the surface of the moon. The selected mission utilizes a spiral trajectory, characteristic of a low-thrust spacecraft, requiring 8 days for a lunar rendezvous and an additional 9 days for return. The ship's configuration consists primarily of an optical train, two hydrogen plasma engines, a 37.1 m box beam truss, a payload module, and fuel tanks. The total mass of the vehicle fully loaded is 63300 kg. A single plasma, regeneratively cooled engine design is incorporated into the two 500 N engines. These are connected to the spacecraft by turntables which allow the vehicle to thrust tangentially to the flight path. Proper collection and transmission of the laser beam to the thrust chambers is provided through the optical train. This system consists of the 23 m diameter primary mirror, a convex parabolic secondary mirror, a beam splitter and two concave parabolic tertiary mirrors. The payload bay is capable of carrying 18000 kg of cargo. The module is located opposite the primary mirror on the main truss. Fuel tanks carrying a maximum of 35000 kg of liquid hydrogen are fastened to tracks which allow the tanks to be moved perpendicular to the main truss. This capability is required to prevent the center of mass from moving out of the thrust vector line. The laser beam is located and tracked by means of an acquisition, pointing and tracking system which can be locked onto the space-based laser station. Correct orientation of the spacecraft with the laser beam is maintained by control moment gyros and reaction control rockets. Additionally an aerobrake configuration was designed to provide the option of using the atmospheric drag in place of propulsion for a return trajectory.

Nuclear Thermal Propulsion Development Risks

NASA Technical Reports Server (NTRS)

Kim, Tony

2015-01-01

There are clear advantages of development of a Nuclear Thermal Propulsion (NTP) for a crewed mission to Mars. NTP for in-space propulsion enables more ambitious space missions by providing high thrust at high specific impulse ((is) approximately 900 sec) that is 2 times the best theoretical performance possible for chemical rockets. Missions can be optimized for maximum payload capability to take more payload with reduced total mass to orbit; saving cost on reduction of the number of launch vehicles needed. Or missions can be optimized to minimize trip time significantly to reduce the deep space radiation exposure to the crew. NTR propulsion technology is a game changer for space exploration to Mars and beyond. However, 'NUCLEAR' is a word that is feared and vilified by some groups and the hostility towards development of any nuclear systems can meet great opposition by the public as well as from national leaders and people in authority. The public often associates the 'nuclear' word with weapons of mass destruction. The development NTP is at risk due to unwarranted public fears and clear honest communication of nuclear safety will be critical to the success of the development of the NTP technology. Reducing cost to NTP development is critical to its acceptance and funding. In the past, highly inflated cost estimates of a full-scale development nuclear engine due to Category I nuclear security requirements and costly regulatory requirements have put the NTP technology as a low priority. Innovative approaches utilizing low enriched uranium (LEU). Even though NTP can be a small source of radiation to the crew, NTP can facilitate significant reduction of crew exposure to solar and cosmic radiation by reducing trip times by 3-4 months. Current Human Mars Mission (HMM) trajectories with conventional propulsion systems and fuel-efficient transfer orbits exceed astronaut radiation exposure limits. Utilizing extra propellant from one additional SLS launch and available energy in the NTP fuel, HMM radiation exposure can be reduced significantly.

Multiple Payload Ejector for Education, Science and Technology Experiments

NASA Technical Reports Server (NTRS)

Lechworth, Gary

2005-01-01

The education research community no longer has a means of being manifested on Space Shuttle flights, and small orbital payload carriers must be flown as secondary payloads on ELV flights, as their launch schedule, secondary payload volume and mass permits. This has resulted in a backlog of small payloads, schedule and cost problems, and an inability for the small payloads community to achieve routine, low-cost access to orbit. This paper will discuss Goddard's Wallops Flight Facility funded effort to leverage its core competencies in small payloads, sounding rockets, balloons and range services to develop a low cost, multiple payload ejector (MPE) carrier for orbital experiments. The goal of the MPE is to provide a low-cost carrier intended primarily for educational flight research experiments. MPE can also be used by academia and industry for science, technology development and Exploration experiments. The MPE carrier will take advantage of the DARPAI NASA partnership to perform flight testing of DARPA s Falcon small, demonstration launch vehicle. The Falcon is similar to MPE fiom the standpoint of focusing on a low-cost, responsive system. Therefore, MPE and Falcon complement each other for the desired long-term goal of providing the small payloads community with a low-cost ride to orbit. The readiness dates of Falcon and MPE are complementary, also. MPE is being developed and readied for flight within 18 months by a small design team. Currently, MPE is preparing for Critical Design Review in fall 2005, payloads are being manifested on the first mission, and the carrier will be ready for flight on the first Falcon demonstration flight in summer, 2006. The MPE and attached experiments can weigh up to 900 lb. to be compatible with Falcon demonstration vehicle lift capabilities fiom Wallops, and will be delivered to the Falcon demonstration orbit - 100 nautical mile circular altitude.

KSC-08pd2641

NASA Image and Video Library

2008-09-17

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, the payload canister is moved on the floor for loading of the Super Lightweight Interchangeable Carrier, or SLIC. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A will be removed. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies for servicing of the Hubble Space Telescope on the STS-125 mission, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

Fuzzy crane control with sensorless payload deflection feedback for vibration reduction

NASA Astrophysics Data System (ADS)

Smoczek, Jaroslaw

2014-05-01

Different types of cranes are widely used for shifting cargoes in building sites, shipping yards, container terminals and many manufacturing segments where the problem of fast and precise transferring a payload suspended on the ropes with oscillations reduction is frequently important to enhance the productivity, efficiency and safety. The paper presents the fuzzy logic-based robust feedback anti-sway control system which can be applicable either with or without a sensor of sway angle of a payload. The discrete-time control approach is based on the fuzzy interpolation of the controllers and crane dynamic model's parameters with respect to the varying rope length and mass of a payload. The iterative procedure combining a pole placement method and interval analysis of closed-loop characteristic polynomial coefficients is proposed to design the robust control scheme. The sensorless anti-sway control application developed with using PAC system with RX3i controller was verified on the laboratory scaled overhead crane.

NASA's Space Launch System: An Evolving Capability for Exploration

NASA Technical Reports Server (NTRS)

Creech, Stephen D.; Robinson, Kimberly F.

2016-01-01

Designed to meet the stringent requirements of human exploration missions into deep space and to Mars, NASA's Space Launch System (SLS) vehicle represents a unique new launch capability opening new opportunities for mission design. NASA is working to identify new ways to use SLS to enable new missions or mission profiles. In its initial Block 1 configuration, capable of launching 70 metric tons (t) to low Earth orbit (LEO), SLS is capable of not only propelling the Orion crew vehicle into cislunar space, but also delivering small satellites to deep space destinations. The evolved configurations of SLS, including both the 105 t Block 1B and the 130 t Block 2, offer opportunities for launching co-manifested payloads and a new class of secondary payloads with the Orion crew vehicle, and also offer the capability to carry 8.4- or 10-m payload fairings, larger than any contemporary launch vehicle, delivering unmatched mass-lift capability, payload volume, and C3.

GAIA payload module mechanical development

NASA Astrophysics Data System (ADS)

Touzeau, S.; Sein, E.; Lebranchu, C.

2017-11-01

Gaia is the European Space Agency's cornerstone mission for global space astrometry. Its goal is to make the largest, most precise three-dimensional map of our Galaxy by surveying an unprecedented number of stars. This paper gives an overview of the mechanical system engineering and verification of the payload module. This development includes several technical challenges. First of all, the very high stability performance as required for the mission is a key driver for the design, which incurs a high degree of stability. This is achieved through the extensive use of Silicon Carbide (Boostec® SiC) for both structures and mirrors, a high mechanical and thermal decoupling between payload and service modules, and the use of high-performance engineering tools. Compliance of payload mass and volume with launcher capability is another key challenge, as well as the development and manufacturing of the 3.2-meter diameter toroidal primary structure. The spacecraft mechanical verification follows an innovative approach, with direct testing on the flight model, without any dedicated structural model.

The Cosmic Ray Energetics And Mass Project

NASA Astrophysics Data System (ADS)

Seo, Eun-Suk; Iss-Cream Collaboration

2017-01-01

The balloon-borne Cosmic Ray Energetics And Mass (CREAM) experiment was flown for 161 days in six flights over Antarctica, the longest known exposure for a single balloon project. Elemental spectra were measured for Z = 1- 26 nuclei over a wide energy range from 1010 to >1014 eV. Building on the success of those balloon flights, one of the two balloon payloads was transformed for exposure on the International Space Station (ISS) Japanese Experiment Module Exposed Facility (JEM-EF). This ISS-CREAM instrument is configured with redundant and complementary particle detectors. The four layers of its finely segmented Silicon Charge Detector provide precise charge measurements, and its ionization calorimeter provides energy measurements. In addition, scintillator-based Top and Bottom Counting Detectors and the Boronated Scintillator Detector distinguish electrons from nuclei. An order of magnitude increase in data collecting power is expected to reach the highest energies practical with direct measurements. Following completion of its qualification tests at NASA Goddard Space Flight Center, the ISS-CREAM payload was delivered to NASA Kennedy Space Center in August 2015 to await its launch to the ISS. While waiting for ISS-CREAM to launch, the other balloon payload including a Transition Radiation Detector, which is too large for the JEM-EF envelope, has been prepared for another Antarctic balloon flight in 2016. This so-called Boron And Carbon Cosmic rays in the Upper Stratosphere (BACCUS) payload will investigate cosmic ray propagation history. The overall project status and future plans will be presented.

Mars MetNet Precursor Mission Status

NASA Astrophysics Data System (ADS)

Harri, Ari-Matti; Aleksashkin, Sergey; Guerrero, Héctor; Schmidt, Walter; Genzer, Maria; Vazquez, Luis; Haukka, Harri

2013-04-01

A new kind of planetary exploration mission for Mars is being developed in collaboration between the Finnish Meteorological Institute (FMI), Lavochkin Association (LA), Space Research Institute (IKI) and Institutio Nacional de Tecnica Aerospacial (INTA). The Mars MetNet mission is based on a new semi-hard landing vehicle called MetNet Lander (MNL), using an inflatable entry and descent system instead of rigid heat shields and parachutes as earlier semi-hard landing devices have used. This way the ratio of the payload mass to the overall mass is optimized. The landing impact will burrow the payload container into the Martian soil providing a more favorable thermal environment for the electronics and a suitable orientation of the telescopic boom with external sensors and the radio link antenna. It is planned to deploy several tens of MNLs on the Martian surface operating at least partly at the same time to allow meteorological network science. For the precursor mission (MMPM) intended to verify the landing concept and key technology during a real Mars mission all qualification activities are completed and the payload and system flight model components are being manufactured. The descent processes dynamic properties are monitored by a special 3-axis accelerometer combined with a 3-axis gyrometer. The data will be sent via auxiliary beacon antenna throughout the descent phase starting shortly after separation from the spacecraft. Details of the current MMPM system and payload configuration and their performance parameters will be shown.

Structures-propulsion interactions and requirements. [large space structures

NASA Technical Reports Server (NTRS)

Coyner, J. V.

1982-01-01

The effects of low-thrust primary propulsion system characteristics on the mass, area, and orbit transfer characteristics of large space systems (LSS) were determined. Three general structural classes of LSS were considered, each with a broad range of diameters and nonstructural surface densities. While transferring the deployed structure from LEO and to GEO, an acceleration range of 0.02 to 0.1 g's was found to maximize deliverable payload based on structural mass impact. After propulsion system parametric analyses considering four propellant combinations produced values for available payload mass, length and volume, a thrust level range which maximizes deliverable LSS diameter was determined corresponding to a structure and propulsion vehicle. The engine start and/or shutdown thrust transients on the last orbit transfer (apogee) burn can impose transient loads which would be greater than the steady-state loads at the burnout acceleration. The effect of the engine thrust transients on the LSS was determined from the dynamic models upon which various engine ramps were imposed.

Aerocapture Guidance and Performance at Mars for High-Mass Systems

NASA Technical Reports Server (NTRS)

Zumwalt, Carlie H.; Sostaric, Ronald r.; Westhelle, Carlos H.; Cianciolo, Alicia Dwyer

2010-01-01

The objective of this study is to understand the performance associated with using the aerocapture maneuver to slow high-mass systems from an Earth-approach trajectory into orbit around Mars. This work is done in conjunction with the Mars Entry Descent and Landing Systems Analysis (EDL-SA) task to explore candidate technologies necessary for development in order to land large-scale payloads on the surface of Mars. Among the technologies considered include hypersonic inflatable aerodynamic decelerators (HIADs) and rigid mid-lift to drag (L/D) aeroshells. Nominal aerocapture trajectories were developed for the mid-L/D aeroshell and two sizes of HIADs, and Monte Carlo analysis was completed to understand sensitivities to dispersions. Additionally, a study was completed in order to determine the size of the larger of the two HIADs which would maintain design constraints on peak heat rate and diameter. Results show that each of the three aeroshell designs studied is a viable option for landing high-mass payloads as none of the three exceed performance requirements.

Integration Of Launch Vehicle Simulation/Analysis Tools And Lunar Cargo Lander Design. Part 2/2

NASA Technical Reports Server (NTRS)

DeJean, George Brian; Shiue, Yeu-Sheng Paul; King, Jeffrey

2005-01-01

Part 2, which will be discussed in this report, will discuss the development of a Lunar Cargo Lander (unmanned launch vehicle) that will transport usable payload from Trans- Lunar Injection to the moon. The Delta IV-Heavy was originally used to transport the Lunar Cargo Lander to TLI, but other launch vehicles have been studied. In order to uncover how much payload is possible to land on the moon, research was needed in order to design the sub-systems of the spacecraft. The report will discuss and compare the use of a hypergolic and cryogenic system for its main propulsion system. The guidance, navigation, control, telecommunications, thermal, propulsion, structure, mechanisms, landing gear, command, data handling, and electrical power sub-systems were designed by scaling off other flown orbiters and moon landers. Once all data was collected, an excel spreadsheet was created to accurately calculate the usable payload that will land on the moon along with detailed mass and volume estimating relations. As designed, The Lunar Cargo Lander can plant 5,400 lbm of usable payload on the moon using a hypergolic system and 7,400 lbm of usable payload on the moon using a cryogenic system.

Catalog of lunar and Mars science payloads

NASA Technical Reports Server (NTRS)

Budden, Nancy Ann (Editor)

1994-01-01

This catalog collects and describes science payloads considered for future robotic and human exploration missions to the Moon and Mars. The science disciplines included are geosciences, meteorology, space physics, astronomy and astrophysics, life sciences, in-situ resource utilization, and robotic science. Science payload data is helpful for mission scientists and engineers developing reference architectures and detailed descriptions of mission organizations. One early step in advanced planning is formulating the science questions for each mission and identifying the instrumentation required to address these questions. The next critical element is to establish and quantify the supporting infrastructure required to deliver, emplace, operate, and maintain the science experiments with human crews or robots. This requires a comprehensive collection of up-to-date science payload information--hence the birth of this catalog. Divided into lunar and Mars sections, the catalog describes the physical characteristics of science instruments in terms of mass, volume, power and data requirements, mode of deployment and operation, maintenance needs, and technological readiness. It includes descriptions of science payloads for specific missions that have been studied in the last two years: the Scout Program, the Artemis Program, the First Lunar Outpost, and the Mars Exploration Program.

Ares V: Game Changer for National Security Launch

NASA Technical Reports Server (NTRS)

Sumrall, Phil; Morris, Bruce

2009-01-01

NASA is designing the Ares V cargo launch vehicle to vastly expand exploration of the Moon begun in the Apollo program and enable the exploration of Mars and beyond. As the largest launcher in history, Ares V also represents a national asset offering unprecedented opportunities for new science, national security, and commercial missions of unmatched size and scope. The Ares V is the heavy-lift component of NASA's dual-launch architecture that will replace the current space shuttle fleet, complete the International Space Station, and establish a permanent human presence on the Moon as a stepping-stone to destinations beyond. During extensive independent and internal architecture and vehicle trade studies as part of the Exploration Systems Architecture Study (ESAS), NASA selected the Ares I crew launch vehicle and the Ares V to support future exploration. The smaller Ares I will launch the Orion crew exploration vehicle with four to six astronauts into orbit. The Ares V is designed to carry the Altair lunar lander into orbit, rendezvous with Orion, and send the mated spacecraft toward lunar orbit. The Ares V will be the largest and most powerful launch vehicle in history, providing unprecedented payload mass and volume to establish a permanent lunar outpost and explore significantly more of the lunar surface than was done during the Apollo missions. The Ares V consists of a Core Stage, two Reusable Solid Rocket Boosters (RSRBs), Earth Departure Stage (EDS), and a payload shroud. For lunar missions, the shroud would cover the Lunar Surface Access Module (LSAM). The Ares V Core Stage is 33 feet in diameter and 212 feet in length, making it the largest rocket stage ever built. It is the same diameter as the Saturn V first stage, the S-IC. However, its length is about the same as the combined length of the Saturn V first and second stages. The Core Stage uses a cluster of five Pratt & Whitney Rocketdyne RS-68B rocket engines, each supplying about 700,000 pounds of thrust. Its propellants are liquid hydrogen and liquid oxygen. The two solid rocket boosters provide about 3.5 million pounds of thrust at liftoff. These 5.5-segment boosters are derived from the 4-segment boosters now used on the Space Shuttle, and are similar to those used in the Ares I first stage. The EDS is powered by one J-2X engine. The J-2X, which has roughly 294,000 pounds of thrust, also powers the Ares I Upper Stage. It is derived from the J-2 that powered the Saturn V second and third stages. The EDS performs two functions. Its initial suborbital burns will place the lunar lander into a stable Earth orbit. After the Orion crew vehicle, launched separately on an Ares I, docks with the lander/EDS stack, EDS will ignite a second time to put the combined 65-metric ton vehicle into a lunar transfer orbit. When it stands on the launch pad at Kennedy Space Center late in the next decade, the Ares V stack will be approximately 381 feet tall and have a gross liftoff mass of 8.1 million pounds. The current point-of-departure design exceeds Saturn V s mass capability by approximately 40 percent. Using the current payload shroud design, Ares V can carry 315,000 pounds to 29-degree low Earth orbit (LEO) or 77,000 pounds to a geosynchronous orbit. Another unique aspect of the Ares V is the 33-foot-diameter payload shroud, which encloses approximately 30,400 cubic feet of usable volume. A larger hypothetical shroud for encapsulating larger payloads has been studied. While Ares V makes possible larger payload masses and volumes, it may alternately make possible more cost-effective mission design if the relevant payload communities are willing to consider an alternative to the existing approach that has driven them to employ complexity to solve current launch vehicle mass and volume constraints. By using Ares V s mass and volume capabilities as margin, payload designers stand to reduce development risk and cost. Significant progress has been made on the Ares V to support a plaed fiscal 2011 authority-to-proceed (ATP) milestone. The Ares V team is actively reaching out to external organizations during this early concept phase to ensure that the Ares V vehicle can be leveraged for national security, science, and commercial development needs. This presentation will discuss Ares V vehicle configuration, the path to the current concept, accomplishments to date, and potential payload utilization opportunities.

Measurement of Atomic Oxygen in Diffuse Aurora and Ion Density in the E-Region

NASA Technical Reports Server (NTRS)

Sharp, William E.

1997-01-01

An ion mass spectrometer (IMS) was refurbished, calibrated and supplied to the University of Colorado payload (Dr. Charles Barth, P.I.) which was launched from White Sands in September of 1993 as NASA 33.062. The nose cone failed to deploy and their were problems with the ACS so the mission was declared a failure. However, the door covering the IMS deployed and the instrument obtained data. The launch occurred shortly after a payload carrying solar x-ray detectors was launched. Thus a small portion of the Colorado payload science was salvaged; namely, the NO(+)/O2(+) ratio to compare with the measured x-ray flux. Figure I shows the NO(+) to O2(+) ratio vs. altitude. The behavior is typical of the E-region.

MINOTAUR (Maryland's innovative orbital technologically advanced University rocket)

NASA Technical Reports Server (NTRS)

Lewis, Mark J.; Akin, Dave; Lind, Charles; Rice, T. (Editor); Vincent, W. (Editor)

1992-01-01

Over the past decade, there has been an increasing interest in designing small commercial launch vehicles. Some of these designs include OSC's Pegasus, and AMROC's Aquila. Even though these vehicles are very different in their overall design characteristics, they all share a common thread of being expensive to design and manufacture. Each of these vehicles has an estimated production and operations cost of over $15000/kg of payload. In response to this high cost factor, the University of Maryland is developing a cost-effective alternative launch vehicle, Maryland's Innovative Orbital Technologically Advanced University Rocket (MINOTAUR). A preliminary cost analysis projects that MINOTAUR will cost under $10000/kg of payload. MINOTAUR will also serve as an enriching project devoted to an entirely student-designed-and-developed launch vehicle. This preliminary design of MINOTAUR was developed entirely by undergraduates in the University of Maryland's Space Vehicle Design class. At the start of the project, certain requirements and priorities were established as a basis from which to begin the design phase: (1) carry a 100 kg payload into a 200 km circular orbit; (2) provide maximum student involvement in the design, manufacturing, and launch phases of the project; and (3) use hybrid propulsion throughout. The following is the list of the project's design priorities (from highest to lowest): (1) safety, (2) cost, (3) minimum development time, (4) maximum use of the off-the-shelf components, (5) performance, and (6) minimum use of pyrotechnics.

Determination of balloon gas mass and revised estimates of drag and virtual mass coefficients

NASA Technical Reports Server (NTRS)

Robbins, E.; Martone, M.

1993-01-01

In support of the NASA Balloon Program, small-scale balloons were flown with varying lifting gas and total system mass. Instrument packages were developed to measure and record acceleration and temperature data during these tests. Top fitting and instrument payload accelerations were measured from launch to steady state ascent and through ballast drop transients. The development of the small lightweight self-powered Stowaway Special instrument packages is discussed along with mathematical models developed to determine gas mass, drag and virtual mass coefficients.

KSC00pp1688

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- The payload transport canister (right) and workers wait for the arrival of the P6 integrated truss segment (left) carried by the overhead crane. After being placed in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1691

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- In the Space Station Processing Facility, the P6 integrated truss segment is placed in the payload transport canister while workers watch its progress. After being secured in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC00pp1691

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- In the Space Station Processing Facility, the P6 integrated truss segment is placed in the payload transport canister while workers watch its progress. After being secured in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1688

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- The payload transport canister (right) and workers wait for the arrival of the P6 integrated truss segment (left) carried by the overhead crane. After being placed in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

Space Launch System Co-Manifested Payload Options for Habitation

NASA Technical Reports Server (NTRS)

Smitherman, David

2015-01-01

The Space Launch System (SLS) has a co-manifested payload capability that will grow over time as the rocket matures and planned upgrades are implemented. The final configuration is planned to be capable of inserting a payload greater than 10 metric tons (mt) into a trans-lunar injection trajectory along with the crew in the Orion capsule and the service module. The co-manifested payload is located below the Orion and its service module in a 10-meter high fairing similar to the way the Saturn launch vehicle carried the lunar lander below the Apollo command and service modules. A variety of approaches have been explored that utilizes this co-manifested payload capability to build up infrastructure in deep space in support of future asteroid, lunar, and Mars mission scenarios. This paper is a report on the findings from the Advanced Concepts Office study team at the NASA Marshall Space Flight Center, working with the Advanced Exploration Systems Program on the Exploration Augmentation Module Project. It includes some of the possible options for habitation in the co-manifested payload volume on SLS. Findings include module designs that can be developed in 10mt increments to support these missions, including overall conceptual layouts, mass properties, and approaches for integration into various scenarios for near-term support of deep space habitat research and technology development, support to asteroid exploration, and long range support for Mars transfer flights.

Mars surface penetrator: System description

NASA Technical Reports Server (NTRS)

Manning, L. A. (Editor)

1977-01-01

A point design of a penetrator system for a Mars mission is described. A strawman payload which is to conduct measurements of geophysical and meteorological parameters is included in the design. The subsystems used in the point design are delineated in terms of power, mass, volume, data, and functional modes. The prospects for survival of the rigors of emplacement are described. Data handling and communications plans are presented to allow consideration of the requirements placed by the penetrator on the orbiter and ground operations. The point design is technically feasible and the payload selection scientifically desirable.

Nuclear electric propulsion mission engineering study. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

1973-01-01

Results of a mission engineering analysis of nuclear-thermionic electric propulsion spacecraft for unmanned interplanetary and geocentric missions are summarized. Critical technologies associated with the development of nuclear electric propulsion (NEP) are assessed. Outer planet and comet rendezvous mission analysis, NEP stage design for geocentric and interplanetary missions, NEP system development cost and unit costs, and technology requirements for NEP stage development are studied. The NEP stage design provides both inherent reliability and high payload mass capability. The NEP stage and payload integration was found to be compatible with the space shuttle.

NASA's Space Launch System: An Evolving Capability for Exploration

NASA Technical Reports Server (NTRS)

Robinson, Kimberly F.; Hefner, Keith; Hitt, David

2015-01-01

Designed to enable human space exploration missions, including eventually landings on Mars, NASA's Space Launch System (SLS) represents a unique launch capability with a wide range of utilization opportunities, from delivering habitation systems into the lunar vicinity to high-energy transits through the outer solar system. The vehicle will be able to deliver greater mass to orbit than any contemporary launch vehicle. SLS will also be able to carry larger payload fairings than any contemporary launch vehicle, and will offer opportunities for co-manifested and secondary payloads.

14 CFR § 125.73 - Contents.

Code of Federal Regulations, 2010 CFR

2017-01-01

...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS... manual must include— (a) The name of each management person who is authorized to act for the certificate... the attention of the pilot in command before, during, and after completion of a flight; (g) Procedures...

14 CFR 125.175 - Protection of other airplane components against fire.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Protection of other airplane components... CERTIFICATION AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD... Requirements § 125.175 Protection of other airplane components against fire. (a) Except as provided in...

14 CFR 125.93 - Airplane limitations.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93 Airplane...

14 CFR 125.355 - Airplane equipment.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Airplane equipment. 125.355 Section 125.355...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Flight Release Rules § 125.355 Airplane...

14 CFR 125.355 - Airplane equipment.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Airplane equipment. 125.355 Section 125.355...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Flight Release Rules § 125.355 Airplane...

14 CFR 125.355 - Airplane equipment.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Airplane equipment. 125.355 Section 125.355...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Flight Release Rules § 125.355 Airplane...

14 CFR 125.93 - Airplane limitations.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93 Airplane...

14 CFR 125.175 - Protection of other airplane components against fire.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Protection of other airplane components... CERTIFICATION AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD... Requirements § 125.175 Protection of other airplane components against fire. (a) Except as provided in...

14 CFR 125.355 - Airplane equipment.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Airplane equipment. 125.355 Section 125.355...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Flight Release Rules § 125.355 Airplane...

14 CFR 125.93 - Airplane limitations.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93 Airplane...

14 CFR 125.175 - Protection of other airplane components against fire.

Code of Federal Regulations, 2012 CFR

2012-01-01

... 14 Aeronautics and Space 3 2012-01-01 2012-01-01 false Protection of other airplane components... CERTIFICATION AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD... Requirements § 125.175 Protection of other airplane components against fire. (a) Except as provided in...

14 CFR 125.175 - Protection of other airplane components against fire.

Code of Federal Regulations, 2011 CFR

2011-01-01

... 14 Aeronautics and Space 3 2011-01-01 2011-01-01 false Protection of other airplane components... CERTIFICATION AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD... Requirements § 125.175 Protection of other airplane components against fire. (a) Except as provided in...

14 CFR 125.75 - Airplane flight manual.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 3 2014-01-01 2014-01-01 false Airplane flight manual. 125.75 Section 125... OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6... Airplane flight manual. (a) Each certificate holder shall keep a current approved Airplane Flight Manual or...

14 CFR 125.93 - Airplane limitations.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 3 2013-01-01 2013-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93 Airplane...

14 CFR 125.73 - Contents.

Code of Federal Regulations, 2010 CFR

1999-01-01

... 14 Aeronautics and Space 2 1999-01-01 1999-01-01 false Contents. 125.73 Section 125.73 AIR CARRIERS AND OPERATORS FOR COMPENSATION OR HIRE: CERTIFICATION AND OPERATIONS CERTIFICATION AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE Special Federal Aviation...

14 CFR 125.175 - Protection of other airplane components against fire.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Protection of other airplane components... CERTIFICATION AND OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD... Requirements § 125.175 Protection of other airplane components against fire. (a) Except as provided in...

14 CFR 125.93 - Airplane limitations.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane limitations. 125.93 Section 125.93...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Airplane Requirements § 125.93 Airplane...

14 CFR 125.355 - Airplane equipment.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane equipment. 125.355 Section 125.355...: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6,000 POUNDS OR MORE; AND RULES GOVERNING PERSONS ON BOARD SUCH AIRCRAFT Flight Release Rules § 125.355 Airplane...

14 CFR 125.75 - Airplane flight manual.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 3 2010-01-01 2010-01-01 false Airplane flight manual. 125.75 Section 125... OPERATIONS: AIRPLANES HAVING A SEATING CAPACITY OF 20 OR MORE PASSENGERS OR A MAXIMUM PAYLOAD CAPACITY OF 6... Airplane flight manual. (a) Each certificate holder shall keep a current approved Airplane Flight Manual or...

Flight Loads and Environments Initiative

NASA Technical Reports Server (NTRS)

Kaufman, Daniel; Kern, Dennis

2005-01-01

A viewgraph presentation on the design of a lightweight non-intrusive force measurement device (FMD) to reduce the cost per effective payload (PL) mass into orbit (CPMO) by improving launch vehicle (LV) loads and environments.

Mars MetNet Mission - Martian Atmospheric Observational Post Network

NASA Astrophysics Data System (ADS)

Hari, Ari-Matti; Haukka, Harri; Aleksashkin, Sergey; Arruego, Ignacio; Schmidt, Walter; Genzer, Maria; Vazquez, Luis; Siikonen, Timo; Palin, Matti

2017-04-01

A new kind of planetary exploration mission for Mars is under development in collaboration between the Finnish Meteorological Institute (FMI), Lavochkin Association (LA), Space Research Institute (IKI) and Institutio Nacional de Tecnica Aerospacial (INTA). The Mars MetNet mission is based on a new semi-hard landing vehicle called MetNet Lander (MNL). The scientific payload of the Mars MetNet Precursor [1] mission is divided into three categories: Atmospheric instruments, Optical devices and Composition and structure devices. Each of the payload instruments will provide significant insights in to the Martian atmospheric behavior. The key technologies of the MetNet Lander have been qualified and the electrical qualification model (EQM) of the payload bay has been built and successfully tested. 1. MetNet Lander The MetNet landing vehicles are using an inflatable entry and descent system instead of rigid heat shields and parachutes as earlier semi-hard landing devices have used. This way the ratio of the payload mass to the overall mass is optimized. The landing impact will burrow the payload container into the Martian soil providing a more favorable thermal environment for the electronics and a suitable orientation of the telescopic boom with external sensors and the radio link antenna. It is planned to deploy several tens of MNLs on the Martian surface operating at least partly at the same time to allow meteorological network science. 2. Strawman Scientific Payload The strawman payload of the two MNL precursor models includes the following instruments: Atmospheric instruments: - MetBaro Pressure device - MetHumi Humidity device - MetTemp Temperature sensors Optical devices: - PanCam Panoramic - MetSIS Solar irradiance sensor with OWLS optical wireless system for data transfer - DS Dust sensor Composition and Structure Devices: Tri-axial magnetometer MOURA Tri-axial System Accelerometer The descent processes dynamic properties are monitored by a special 3-axis accelerometer combined with a 3-axis gyrometer. The data will be sent via auxiliary beacon antenna throughout the descent phase starting shortly after separation from the spacecraft. MetNet Mission payload instruments are specially designed to operate under very low power conditions. MNL flexible solar panels provides a total of approximately 0.7-0.8 W of electric power during the daylight time. As the provided power output is insufficient to operate all instruments simultaneously they are activated sequentially according to a specially designed cyclogram table which adapts itself to the different environmental constraints. 3. Mission Status he eventual goal is to create a network of atmospheric observational posts around the Martian surface. Even if the MetNet mission is focused on the atmospheric science, the mission payload will also include additional kinds of geophysical instrumentation. The next step is the MetNet Precursor Mission that will demonstrate the technical robustness and scientific capabilities of the MetNet type of landing vehicle. Definition of the Precursor Mission and discussions on launch opportunities are currently under way. The first MetNet Science Payload Precursors have already been successfully completed, e,g, the REMS/MSL and DREAMS/Exomars-2016. The next MetNet Payload Precursors will be METEO/Exomars-2018 and MEDA/Mars-2020. The baseline program development funding exists for the next seven years. Flight unit manufacture of the payload bay takes about 18 months, and it will be commenced after the Precursor Mission has been defined. References [1] http://metnet.fmi.fi

Experimental Study of Convective Heating on the Back Face and Payload of a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Aeroshell

NASA Technical Reports Server (NTRS)

Hollis, Brian R.; Berry, Scott A.; Hollingsworth, Kevin E.; Wright, Sheila A.

2017-01-01

A wind tunnel test program has been conducted to define convective heating environments on the back-face of a Hypersonic Inflatable Aerodynamic Decelerator aeroshell. Wind tunnel testing was conducted at Mach 6 and Mach 10 at unit Reynolds numbers from 0.5×10(exp 6)/ft to 3.9×10(exp 6)/ft on a 6.3088 in diameter aeroshell model. Global heating data were obtained through phosphor thermography on the aeroshell back face, as well as on the payload and the aeroshell front face. For all test conditions, laminar flow was produced on the aeroshell front face, while the separated wake shear layer and aeroshell back-face boundary layer were transitional or turbulent. Along the leeward centerline of the aeroshell back face and payload centerbody, heating levels increased with both free stream Reynolds number and angle of attack. The Reynolds number dependency was due to increasing strength of wake turbulence with Reynolds number. The angle-of-attack dependency was due to movement of the wake-vortex reattachment point on the aeroshell back face. The maximum heating levels on the aeroshell back face and payload were approximately 5% to 6%, respectively, of the aeroshell front-face stagnation point. To allow for extrapolation of the ground test data to flight conditions, the back face and payload heating levels were correlated as a function of aeroshell front-face peak momentum thickness Reynolds numbers.

Project Aether Aurora: STEM outreach near the arctic circle

NASA Astrophysics Data System (ADS)

Longmier, B. W.; Bering, E. A.

2012-12-01

Project Aether is a program designed to immerse high-school through graduate students to field research in some of the fields of STEM. The program leaders launch high altitude weather balloons in collaboration with schools and students to teach physics concepts, experimental research skills, and to make space exploration accessible to students. A weather balloon lifts a specially designed payload package that is composed of HD cameras, GPS tracking devices, and other science equipment. The payload is constructed and attached to the balloon by the students with low-cost materials. The balloon and payload are launched with FAA clearance from a site chosen based on wind patterns and predicted landing locations. The balloon ascends over 2 hours to a maximum altitude of 100,000 feet where it bursts and allows the payload to slowly descend using a built-in parachute. The balloon's location is monitored during its flight by GPS-satellite relay. Most of the science and video data are recorded on SD cards using an Arduino digitizer. The payload is located using the GPS device. The science data are recovered from the payload and shared with the students. In April 2012, Project Aether leaders conducted a field campaign near Fairbanks Alaska, sending several student-built experiments to an altitude of 30km, underneath several strong auroral displays. Auroral physics experiments that can be done on ultra small balloons (5 cubic meters) include electric field and magnetic fluctuation observations, full spectrum and narrow band optical imaging, GPS monitoring of the total electron content of the ionosphere, x-ray detection and infrared and UV spectroscopy. The actual undergraduate student experiments will be reviewed and some data presented.; Balloon deployment underneath aurora, Fairbanks Alaska, 2012.

Design, Development, and Integration of A Space Shuttle Orbiter Bay 13 Payload Carrier

NASA Technical Reports Server (NTRS)

Spencer, Susan H.; Phillips, Michael W.; Upton, Lanny (Technical Monitor)

2002-01-01

Bay 13 of the Space Shuttle Orbiter has been limited to small sidewall mounted payloads and ballast. In order to efficiently utilize this space, a concept was developed for a cross-bay cargo carrier to mount Orbital Replacement Units (ORU's) for delivery to the International Space Station and provide additional opportunities for science payloads, while meeting the Orbiter ballast requirements. The Lightweight Multi-Purpose Experiment Support Structure (MPESS) Carrie (LMC) was developed and tested by NASA's Marshall Space Flight Center and the Boeing Company. The Multi-Purpose Experiment Support Structure (MPESS), which was developed for the Spacelab program was modified, removing the keel structure and relocating the sill trunnions to fit in Bay 13. Without the keel fitting, the LMC required a new and innovative concept for transferring Y loads into the Orbiter structure. Since there is no keel fitting available in the Bay 13 location, the design had to utilize the longeron bridge T-rail to distribute the Y loads. This concept has not previously been used in designing Shuttle payloads. A concept was developed to protect for Launch-On-Need ORU's, while providing opportunities for science payloads. Categories of potential ORU's were defined, and Get-Away Special (GAS) payloads of similar mass properties were provided by NASA's Goddard Space Flight Center. Four GAS payloads were manifest as the baseline configuration, preserving the capability to swap up to two ORU's for the corresponding science payloads, after installation into the Orbiter cargo bay at the pad, prior to closeout. Multiple configurations were considered for the analytical integration, to protect for all defined combinations of ORU's and GAS payloads. The first physical integration of the LMC war performed by Goddard Space Flight Center and Kennedy Space Center at an off-line facility at Kennedy Space Center. This paper will discuss the design challenges, structural testing, analytical and physical integration for the LMC's successful maiden flight on STS-108/ISS UF-1 mission in December 2001.

Preliminary Structural Design Considerations and Mass Efficiencies for Lunar Surface Manipulator Concepts

NASA Technical Reports Server (NTRS)

Dorsey, John T.; Mikulas, Martin M.; Doggett, William R.

2008-01-01

The mass and sizing characteristics of manipulators for Lunar and Mars planetary surface applications are investigated by analyzing three structural configurations: a simple cantilevered boom with a square tubular cross-section; a hybrid cable/boom configuration with a square tubular cross-section support structure; and a hybrid cable/boom configuration with a square truss cross-section support structure. Design procedures are developed for the three configurations and numerical examples are given. A new set of performance parameters are developed that relate the mass of manipulators and cranes to a loading parameter. These parameters enable the masses of different manipulator configurations to be compared over a wide range of design loads and reach envelopes (radii). The use of these parameters is demonstrated in the form of a structural efficiency chart using the newly considered manipulator configurations. To understand the performance of Lunar and Mars manipulators, the design procedures were exercised on the three manipulator configurations assuming graphite/epoxy materials for the tubes and trusses. It is also assumed that the actuators are electric motor, gear reduction systems. Numerical results for manipulator masses and sizes are presented for a variety of manipulator reach and payload mass capabilities. Results are presented that demonstrate the sensitivity of manipulator mass to operational radius, tip force, and actuator efficiency. The effect of the value of gravitational force on the ratio of manipulator-mass to payload-mass is also shown. Finally, results are presented to demonstrate the relative mass reduction for the use of graphite/epoxy compared to aluminum for the support structure.

ARM Aerial Facility ArcticShark Unmanned Aerial System

NASA Astrophysics Data System (ADS)

Schmid, B.; Hubbell, M.; Mei, F.; Carroll, P.; Mendoza, A.; Ireland, C.; Lewko, K.

2017-12-01

The TigerShark Block 3 XP-AR "ArcticShark" Unmanned Aerial System (UAS), developed and manufactured by Navmar Applied Sciences Corporation (NASC), is a single-prop, 60 hp rotary-engine platform with a wingspan of 6.5 m and Maximum Gross Takeoff Weight of 295 Kg. The ArcticShark is owned by the U.S. Department of Energy (DOE) and has been operated by Pacific Northwest National Laboratory (PNNL) since March 2017. The UAS will serve as an airborne atmospheric research observatory for DOE ARM, and, once fully operational, can be requested through ARM's annual call for proposals. The Arctic Shark is anticipated to measure a wide range of radiative, aerosol, and cloud properties using a variable instrument payload weighing up to 46 Kg. SATCOM-equipped, it is capable of taking measurements up to altitudes of 5.5 Km over ranges of up to 500 Km. The ArcticShark operates at airspeeds of 30 to 40 m/s, making it capable of slow sampling. With a full fuel load, its endurance exceeds 8 hours. The aircraft and its Mobile Operations Center (MOC) have been hardened specifically for operations in colder temperatures.ArcticShark's design facilitates rapid integration of various types of payloads. 2500 W of its 4000 W electrical systems is dedicated to payload servicing. It has an interior payload volume of almost 85 L and four wing-mounted pylons capable of carrying external probes. Its payload bay volume, electrical power, payload capacity, and flight characteristics enable the ArcticShark to accommodate multiple combinations of payloads in numerous configurations. Many instruments will be provided by the ARM Aerial Facility (AAF), but other organizations may eventually propose instrumentation for specific campaigns. AAF-provided measurement capabilities will include the following atmospheric state and thermodynamics: temperature, pressure, winds; gases: H2O and CO2; up- and down-welling broadband infrared and visible radiation; surface temperature; aerosol number concentration, size distribution, absorption composition (filter samples), and cloud-droplet size distribution.

Evaluation of on-board hydrogen storage methods f or high-speed aircraft

NASA Technical Reports Server (NTRS)

Akyurtlu, Ates; Akyurtlu, Jale F.

1991-01-01

Hydrogen is the fuel of choice for hypersonic vehicles. Its main disadvantage is its low liquid and solid density. This increases the vehicle volume and hence the drag losses during atmospheric flight. In addition, the dry mass of the vehicle is larger due to larger vehicle structure and fuel tankage. Therefore it is very desirable to find a fuel system with smaller fuel storage requirements without deteriorating the vehicle performance substantially. To evaluate various candidate fuel systems, they were first screened thermodynamically with respect to their energy content and cooling capacities. To evaluate the vehicle performance with different fuel systems, a simple computer model is developed to compute the vehicle parameters such as the vehicle volume, dry mass, effective specific impulse, and payload capacity. The results indicate that if the payload capacity (or the gross lift-off mass) is the most important criterion, only slush hydrogen and liquid hydrogen - liquid methane gel shows better performance than the liquid hydrogen vehicle. If all the advantages of a smaller vehicle are considered and a more accurate mass analysis can be performed, other systems using endothermic fuels such as cyclohexane, and some boranes may prove to be worthy of further consideration.

High Efficiency Synchronous Rectification in Spacecraft

NASA Technical Reports Server (NTRS)

Krauhamer, S.; Das, R.; Vorperian, V.; White, J.; Bennett, J.; Rogers, D.

1993-01-01

This paper examines the implementaion of MOSFETs as synchronous rectifiers which results in a substantial improvement in power processing efficency and therefore may result in significant reduction of spacecraft mass and volum for the same payload.

CubeSat Material Limits For Design for Demise

NASA Technical Reports Server (NTRS)

Kelley, R. L.; Jarkey, D. R.

2014-01-01

The CubeSat form factor of nano-satellite (a satellite with a mass between one and ten kilograms) has grown in popularity due to their ease of construction and low development and launch costs. In particular, their use as student led payload design projects has increased due to the growing number of launch opportunities. CubeSats are often deployed as secondary or tertiary payloads on most US launch vehicles or they may be deployed from the ISS. The focus of this study will be on CubeSats launched from the ISS. From a space safety standpoint, the development and deployment processes for CubeSats differ significantly from that of most satellites. For large satellites, extensive design reviews and documentation are completed, including assessing requirements associated with reentry survivability. Typical CubeSat missions selected for ISS deployment have a less rigorous review process that may not evaluate aspects beyond overall design feasibility. CubeSat design teams often do not have the resources to ensure their design is compliant with reentry risk requirements. A study was conducted to examine methods to easily identify the maximum amount of a given material that can be used in the construction of a CubeSats without posing harm to persons on the ground. The results demonstrate that there is not a general equation or relationship that can be used for all materials; instead a limiting value must be defined for each unique material. In addition, the specific limits found for a number of generic materials that have been previously used as benchmarking materials for reentry survivability analysis tool comparison will be discussed.

CubeSat Material Limits for Design for Demise

NASA Technical Reports Server (NTRS)

Kelley, R. L.; Jarkey, D. R.

2014-01-01

The CubeSat form factor of nano-satellite (a satellite with a mass between one and ten kilograms) has grown in popularity due to their ease of construction and low development and launch costs. In particular, their use as student led payload design projects has increased due to the growing number of launch opportunities. CubeSats are often deployed as secondary or tertiary payloads on most US launch vehicles or they may be deployed from the ISS. The focus of this study will be on CubeSats launched from the ISS. From a space safety standpoint, the development and deployment processes for CubeSats differ significantly from that of most satellites. For large satellites, extensive design reviews and documentation are completed, including assessing requirements associated with re-entry survivability. Typical CubeSat missions selected for ISS deployment have a less rigorous review process that may not evaluate aspects beyond overall design feasibility. CubeSat design teams often do not have the resources to ensure their design is compliant with re-entry risk requirements. A study was conducted to examine methods to easily identify the maximum amount of a given material that can be used in the construction of a CubeSats without posing harm to persons on the ground. The results demonstrate that there is not a general equation or relationship that can be used for all materials; instead a limiting value must be defined for each unique material. In addition, the specific limits found for a number of generic materials that have been previously used as benchmarking materials for re-entry survivability analysis tool comparison will be discussed.

Benefits of in situ propellant utilization for a Mars sample return mission

NASA Technical Reports Server (NTRS)

Wadel, Mary F.

1993-01-01

Previous Mars rover sample return mission studies have shown a requirement for Titan 4 or STS Space Shuttle launch vehicles to complete a sample return from a single Mars site. These studies have either used terrestrial propellants or considered in situ production of methane and oxygen for the return portion of the mission. Using in situ propellants for the return vehicles reduces the Earth launch mass and allows for a smaller Earth launch vehicle, since the return propellant is not carried from Earth. Carbon monoxide and oxygen (CO/O2) and methane and oxygen (CH4/O2) were investigated as in situ propellants for a Mars sample return mission and the results were compared to a baseline study performed by the Jet Propulsion Laboratory using terrestrial propellants. Capability for increased sample return mass, use of an alternate launch vehicle, and an additional mini-rover as payload were included. CO/O2 and CH4/O2 were found to decrease the baseline Earth launch mass by 13.6 and 9.2 percent, respectively. This resulted in higher payload mass margins for the baseline Atlas 2AS launch vehicle. CO/O2 had the highest mass margin. And because of this, it was not only possible to increase the sample return mass and carry an additional mini-rover, but was also possible to use the smaller Atlas 2A launch vehicle.

Opportunities for Geoscience Research Onboard Virgin Galactic's SpaceShipTwo

NASA Astrophysics Data System (ADS)

Pomerantz, W.; Beerer, I.; Stephens, K.; Griffith, J.; Persall, W.; Tizard, J.

2012-12-01

Virgin Galactic has developed a reusable spaceplane, called SpaceShipTwo (SS2), designed to make routine voyages into suborbital space. SS2 is air-launched from a jet aircraft at an altitude of 50,000 ft. before igniting its rocket motor engine. The vehicle reaches a maximum apogee as high as 110 km before gliding to a conventional runway landing. With the ability to fly multiple times per week, SS2 will be capable of providing routine access to a rarely sampled and poorly understood region of the atmosphere and ionosphere, making it a valuable platform for geoscience research. With a payload capacity up to 1300 lbs., SS2 provides access to space and the upper atmosphere for substantially larger payloads than sounding rockets and at a dramatically lower cost than orbital satellites. The main cabin provides as much as 500 cubic ft. of useable volume in a shirt-sleeve environment and payload mounting interfaces that are compatible with standard architectures, such as Middeck Lockers, Cargo Transfer Bags, and server racks. A flight test engineer will be available on board to operate payloads during flight. In the future, SS2 will also offer a variety of external payload mounting locations, enabling researchers to make frequent in situ measurements in the mesosphere (50-90 km), lower thermosphere (above 80 km), and lower ionosphere (above 60 km). SS2 may also offer optical quality windows, allowing optical investigations from main cabin payloads. Researchers will have access to their payloads until just hours before flight and within three hours post-flight. While commercial operations will begin out of Spaceport America in New Mexico, SS2 may eventually be able to launch from a variety of geographic locations. Funding to develop and fly payloads for SS2 is currently available through many NASA programs including the Flight Opportunities Program and the Game Changing Development Program. Virgin Galactic expects the SS2 research platform to enable significant progress in atmospheric chemistry and dynamics, climate science, space weather, numerical weather predictions, and many other fields of geoscience.

UltraSail CubeSat Solar Sail Flight Experiment

NASA Technical Reports Server (NTRS)

Carroll, David; Burton, Rodney; Coverstone, Victoria; Swenson, Gary

2013-01-01

UltraSail is a next-generation, highrisk, high-payoff sail system for the launch, deployment, stabilization, and control of very large (km2 class) solar sails enabling high payload mass fractions for interplanetary and deep space spacecraft. UltraSail is a non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation- flying microsatellites with an innovative solar sail architecture to achieve controllable sail areas approaching 1 km2, sail subsystem area densities approaching 1 g/m2, and thrust levels many times those of ion thrusters used for comparable deep space missions. UltraSail can achieve outer planetary rendezvous, a deep-space capability now reserved for high-mass nuclear and chemical systems. There is a twofold rationale behind the UltraSail concept for advanced solar sail systems. The first is that sail-andboom systems are inherently size-limited. The boom mass must be kept small, and column buckling limits the boom length to a few hundred meters. By eliminating the boom, UltraSail not only offers larger sail area, but also lower areal density, allowing larger payloads and shorter mission transit times. The second rationale for UltraSail is that sail films present deployment handling difficulties as the film thickness approaches one micrometer. The square sail requires that the film be folded in two directions for launch, and similarly unfolded for deployment. The film is stressed at the intersection of two folds, and this stress varies inversely with the film thickness. This stress can cause the film to yield, forming a permanent crease, or worse, to perforate. By rolling the film as UltraSail does, creases are prevented. Because the film is so thin, the roll thickness is small. Dynamic structural analysis of UltraSail coupled with dynamic control analysis shows that the system can be designed to eliminate longitudinal torsional waves created while controlling the pitch of the blades, while using solar photon pressure to slew the spin axis. Vacuum tests have also verified that electrostatic and molecular adhesion forces can substantially be eliminated by making the film electrically conductive, reducing the peel force of the film off the storage roll to levels of 100s of micro-N. The innovation demonstrated the capability of deploying a six-micron aluminum- coated film from a reel through a slit in vacuum. The innovation also demonstrated a spin-stabilized method for deploying a long reel of solar sail film using solar pressure to spin-up and orbit raise the satellite, and also a gravity gradient method for deploying a long reel of solar sail film using solar pressure to orbit raise the satellite. The solar sail mass fraction of 25% is consistent with high specific impulse ion systems, but without the added weight and cost of a power source and processing unit. The large sail area, coupled with low film density, is giving UltraSail a high payload fraction. The UltraSail deployment scheme unrolls a micrometerscale reflection-coated polyimide film from a storage mandrel to a maximum length of several kilometers with the aid of a blade tip satellite.

Prediction of three sigma maximum dispersed density for aerospace applications

NASA Technical Reports Server (NTRS)

Charles, Terri L.; Nitschke, Michael D.

1993-01-01

Free molecular heating (FMH) is caused by the transfer of energy during collisions between the upper atmosphere molecules and a space vehicle. The dispersed free molecular heating on a surface is an important constraint for space vehicle thermal analyses since it can be a significant source of heating. To reduce FMH to a spacecraft, the parking orbit is often designed to a higher altitude at the expense of payload capability. Dispersed FMH is a function of both space vehicle velocity and atmospheric density, however, the space vehicle velocity variations are insignificant when compared to the atmospheric density variations. The density of the upper atmosphere molecules is a function of altitude, but also varies with other environmental factors, such as solar activity, geomagnetic activity, location, and time. A method has been developed to predict three sigma maximum dispersed density for up to 15 years into the future. This method uses a state-of-the-art atmospheric density code, MSIS 86, along with 50 years of solar data, NASA and NOAA solar activity predictions for the next 15 years, and an Aerospace Corporation correlation to account for density code inaccuracies to generate dispersed maximum density ratios denoted as 'K-factors'. The calculated K-factors can be used on a mission unique basis to calculate dispersed density, and hence dispersed free molecular heating rates. These more accurate K-factors can allow lower parking orbit altitudes, resulting in increased payload capability.

Database of proposed payloads and instruments for SEI missions

NASA Technical Reports Server (NTRS)

Barlow, N. G.

1992-01-01

A database of all payloads and instruments proposed for lunar and Mars missions was compiled by the author for the Exploration Programs Office at NASA's Johnson Sapce Center. The database is an outgrowth of the document produced by C. J. Budney et al. at the Jet Propulsion Laboratory in 1991. The present database consists not only of payloads proposed for human exploratory missions of the Moon and Mars, but also experiments selected or proposed for robotic precursor missions such as Lunar Scout, Mars Observer, and MESUR. The database consists of two parts: a written payload description and a matrix that provides a breakdown of payload components. Each payload description consists of the following information: (1) the rationale for why the instrument or payload package is being proposed for operation on the Moon or Mars; (2) a description of how the instrument works; (3) a breakdown of the payload, providing detailed information about the mass, volume, power requirements, and data rates for the constituent pieces of the experiment; (4) estimates of the power consumption and data rate; (5) how the data will be returned to Earth and distributed to the scientific community; (6) any constraints on the location or conditions under which the instrument can or cannot operate; (7) what type of crew interaction (if any) is needed; (8) how the payload is to be delivered to the lunar or martian surface (along with alternative delivery options); (9) how long the instrument or payload package will take to set up; (10) what type of maintenance needs are anticipated for the experiment; (11) stage of development for the instrument and environmental conditions under which the instrument has been tested; (12) an interface required by the instrument with the lander, a rover, an outpost, etc.; (13) information about how often the experiment will need to be resupplied with parts or consumables, if it is to be resupplied; (14) the name and affiliation of a contact person for the experiment; and (15) references where further information about the experiment can be found.

Transient excitation and mechanical admittance test techniques for prediction of payload vibration environments

NASA Technical Reports Server (NTRS)

Kana, D. D.; Vargas, L. M.

1977-01-01

Transient excitation forces were applied separately to simple beam-and-mass launch vehicle and payload models to develop complex admittance functions for the interface and other appropriate points on the structures. These measured admittances were then analytically combined by a matrix representation to obtain a description of the coupled system dynamic characteristics. Response of the payload model to excitation of the launch vehicle model was predicted and compared with results measured on the combined models. These results are also compared with results of earlier work in which a similar procedure was employed except that steady-state sinusoidal excitation techniques were included. It is found that the method employing transient tests produces results that are better overall than the steady state methods. Furthermore, the transient method requires far less time to implement, and provides far better resolution in the data. However, the data acquisition and handling problem is more complex for this method. It is concluded that the transient test and admittance matrix prediction method can be a valuable tool for development of payload vibration tests.

Comparative assessment of anti-sway control strategy for tower crane system

NASA Astrophysics Data System (ADS)

Samin, Reza Ezuan; Mohamed, Zaharuddin

2017-09-01

Tower crane is also known as rotary crane and widely used in constructions due to limited human capability to carry the various types of load at the construction site. In general crane is used for the purpose of loading and unloading heavy material from one place to another. However, in order to transfer the material in minimum time from one location to another, swaying of the payload will occur. Hence, this research presents the investigation of tower crane system which mainly focusing on the swaying angle of the payload by implementing conventional and intelligent controllers. Its mathematical modeling is developed using the Newton's Second Law and simulation is done within the MATLAB/Simulink environment. Simulation results are presented in cart trajectory capability and payload sway angle reduction. A comparative assessment between conventional controller and intelligent controller for the tower crane system are presented and discussed. Furthermore, the effect of various rope length and payload mass of the tower crane system to the performance of trajectory capability and sway angle reduction are also presented and discussed.

Project UM-HAUL (UnManned Heavy pAyload Unloader and Lander): The design of a reusable lunar lander with an independent cargo unloader

NASA Technical Reports Server (NTRS)

1991-01-01

Project UM-Haul is the preliminary design of a reusable lunar transportation vehicle that travels between a lunar parking orbit and the lunar surface. This vehicle is an indispensible link in the overall task of establishing a lunar base as defined by the NASA Space Exploration Initiative. The response to this need consists of two independent vehicles: a lander and an unloader. The system can navigate and unload itself with a minimum amount of human intervention. The design addresses structural analysis, propulsion, power, controls, communications, payload handling and orbital operations. The Lander has the capacity to decend from low lunar orbit (LLO) to the lunar surface carrying a 7000 kg payload, plus the unloader, plus propellant for ascent to LLO. The Lander employs the Unloader by way of a motorized ramp. The Unloader is a terrain vehicle capable of carrying cargoes of 8,500 kg mass and employs a lift system to lower payloads to the ground. The system can perform ten missions before requiring major servicing.

Radioisotope Stirling Engine Powered Airship for Atmospheric and Surface Exploration of Titan

NASA Technical Reports Server (NTRS)

Colozza, Anthony J.; Cataldo, Robert L.

2014-01-01

The feasibility of an advanced Stirling radioisotope generator (ASRG) powered airship for the near surface exploration of Titan was evaluated. The analysis did not consider the complete mission only the operation of the airship within the atmosphere of Titan. The baseline airship utilized two ASRG systems with a total of four general-purpose heat source (GPHS) blocks. Hydrogen gas was used to provide lift. The ASRG systems, airship electronics and controls and the science payload were contained in a payload enclosure. This enclosure was separated into two sections, one for the ASRG systems and the other for the electronics and payload. Each section operated at atmospheric pressure but at different temperatures. The propulsion system consisted of an electric motor driving a propeller. An analysis was set up to size the airship that could operate near the surface of Titan based on the available power from the ASRGs. The atmospheric conditions on Titan were modeled and used in the analysis. The analysis was an iterative process between sizing the airship to carry a specified payload and the power required to operate the electronics, payload and cooling system as well as provide power to the propulsion system to overcome the drag on the airship. A baseline configuration was determined that could meet the power requirements and operate near the Titan surface. From this baseline design additional trades were made to see how other factors affected the design such as the flight altitude and payload mass and volume.

Evaluation of Oxygen Interactions with Materials 3: Mission and induced environments

NASA Technical Reports Server (NTRS)

Koontz, Steven L.; Leger, Lubert J.; Rickman, Steven L.; Hakes, Charles L.; Bui, David T.; Hunton, Donald; Cross, Jon B.

1995-01-01

The Evaluation of Oxygen Interactions with Materials 3 (EOIM-3) flight experiment was developed to obtain benchmark atomic oxygen/material reactivity data. The experiment was conducted during Space Shuttle mission 46 (STS-46), which flew July 31 to August 7, 1992. Quantitative interpretation of the materials reactivity measurements requires a complete and accurate definition of the space environment exposure, including the thermal history of the payload, the solar ultraviolet exposure, the atomic oxygen fluence, and any spacecraft outgassing contamination effects. The thermal history of the payload was measured using twelve thermocouple sensors placed behind selected samples and on the EOIM-3 payload structure. The solar ultraviolet exposure history of the EOIM-3 payload was determined by analysis of the as-flown orbit and vehicle attitude combined with daily average solar ultraviolet and vacuum ultraviolet (UV/VUV) fluxes. The atomic oxygen fluence was assessed in three different ways. First, the O-atom fluence was calculated using a program that incorporates the MSIS-86 atmospheric model, the as-flown Space Shuttle trajectory, and solar activity parameters. Second, the oxygen atom fluence was estimated directly from Kapton film erosion. Third, ambient oxygen atom measurements were made using the quadrupole mass spectrometer on the EOIM-3 payload. Our best estimate of the oxygen atom fluence as of this writing is 2.3 +/- 0.3 x 10(exp 20) atoms/sq cm. Finally, results of post-flight X-ray photoelectron spectroscopy (XPS) surface analyses of selected samples indicate low levels of contamination on the payload surface.

Analysis of a Radioisotope Thermal Rocket Engine

NASA Technical Reports Server (NTRS)

Machado-Rodriguez, Jonathan P.; Landis, Geoffrey A.

2016-01-01

The Triton Hopper is a concept for a global hopper vehicle which uses a radioisotope rocket engine and In-situ propellant acquisition to explore the surface of Neptune's moon, Triton. The current Triton Hopper concept stores heated Nitrogen in a spherical tank to be used as the propellant. The aim of the research was to investigate the benefits of storing propellant at ambient temperature and heating it through the use of a thermal block during engine operation, as opposed to storing gas at a high temperature. Lithium, Lithium Fluoride and Beryllium were considered as possible materials for the thermal block. A heat energy analysis indicated that a lithium thermal mass would provide the highest heat energy for a temperature change from 900 Celsius to -100 Celsius. A heat transfer analysis was performed for Nitrogen at -100 Celsius flowing through 1000 passages inside a 1kg lithium thermal block at a temperature of 900 Celsius. The system was analyzed as turbulent flow through a tube with constant surface temperature. The analysis indicated that the propellant reached a maximum temperature of 877 Celsius before entering the nozzle. At this exit temperature, the average specific impulse [I(sub sp)] of the engine was determined to be 157s. Previous studies for the stored heated gas concept suggest that the engine would have an average I(sub sp) of approximately 52s. Thus, the use of a thermal block concept results in a 200 percent engine performance increase. In addition, a tank sizing study was performed to determine if the concept is feasible in terms of mass requirements. The mass for a spherical carbon fiber COPV storing 35kg of nitrogen at an initial temperature of -100 Celsius and a pressure of 1000psia, was determined to be 7.2kg. The specific impulse analysis indicated that the maximum engine performance is obtained for a mass ratio of 5kg of Nitrogen per every 1kg of lithium thermal mass. Thus for 35kg of Nitrogen the total thermal mass would be 7kg. This brings the total mass of the system to 49.2.kg which is less than the 56kg landing payload capacity of the Triton Hopper. Finally, an insulation analysis using 10mm of MLI insulation indicated that a total of 22 watts of heat are lost to the environment. With the heat loss known, the power required to heat the thermal mass to 900 Celsius in 24 days was determined to be 2.15 watts. The study's results allowed us to conclude that the thermal mass concept is the better option due to the performance increase provided, the low power requirement and its compliance with the landing mass requirement of the Triton Hopper.

Compact Micromachined Bandpass Filters for Infrared Planetary Spectroscopy

NASA Technical Reports Server (NTRS)

Brown, Ari D.; Aslam, Shahid; Chervenak, James A.; Huang, Wei-Chung; Merrell, Willie; Quijada, Manuel

2011-01-01

The thermal instrument strawman payload of the Jupiter Europa Orbiter on the Europa Jupiter Science Mission will map out thermal anomalies, the structure, and atmospheric conditions of Europa and Jupiter within the 7-100 micron spectral range. One key requirement for the payload is that the mass cannot exceed 3.7 kg. Consequently, a new generation of light-weight miniaturized spectrometers needs to be developed. On the path toward developing these spectrometers is development of ancillary miniaturized spectroscopic components. In this paper, we present a strategy for making radiation hard and low mass FIR band pass metal mesh filters. Our strategy involves using MEMS-based fabrication techniques, which will permit the quasi-optical filter structures to be made with micron-scale precision. This will enable us to achieve tight control over both the pass band of the filter and the micromachined silicon support structure architecture, which will facilitate integration of the filters for a variety of applications.

Aerocapture Inflatable Decelerator for Planetary Entry

NASA Technical Reports Server (NTRS)

Reza, Sajjad; Hund, Richard; Kustas, Frank; Willcockson, William; Songer, Jarvis; Brown, Glen

2007-01-01

Forward Attached Inflatable Decelerators, more commonly known as inflatable aeroshells, provide an effective, cost efficient means of decelerating spacecrafts by using atmospheric drag for aerocapture or planetary entry instead of conventional liquid propulsion deceleration systems. Entry into planetary atmospheres results in significant heating and aerodynamic pressures which stress aeroshell systems to their useful limits. Incorporation of lightweight inflatable decelerator surfaces with increased surface-area footprints provides the opportunity to reduce heat flux and induced temperatures, while increasing the payload mass fraction. Furthermore, inflatable aeroshell decelerators provide the needed deceleration at considerably higher altitudes and Mach numbers when compared with conventional rigid aeroshell entry systems. Inflatable aeroshells also provide for stowage in a compact space, with subsequent deployment of a large-area, lightweight heatshield to survive entry heating. Use of a deployable heatshield decelerator enables an increase in the spacecraft payload mass fraction and may eliminate the need for a spacecraft backshell.

Project ECHO: Electronic Communications from Halo Orbit

NASA Technical Reports Server (NTRS)

Borrelli, Jason; Cooley, Bryan; Debole, Marcy; Hrivnak, Lance; Nielsen, Kenneth; Sangmeister, Gary; Wolfe, Matthew

1994-01-01

The design of a communications relay to provide constant access between the Earth and the far side of the Moon is presented. Placement of the relay in a halo orbit about the L2 Earth-Moon Lagrange point allows the satellite to maintain constant simultaneous communication between Earth and scientific payloads on the far side of the Moon. The requirements of NASA's Discovery-class missions adopted and modified for this design are: total project cost should not exceed $150 million excluding launch costs, launch must be provided by Delta-class vehicle, and the satellite should maintain an operational lifetime of 10 to 15 years. The spacecraft will follow a transfer trajectory to the L2 point, after launch by a Delta II 7925 vehicle in 1999. Low-level thrust is used for injection into a stationkeeping-free halo orbit once the spacecraft reaches the L2 point. The shape of this halo orbit is highly elliptical with the maximum excursion from the L2 point being 35000 km. A spun section and despun section connected through a bearing and power transfer assembly (BAPTA) compose the structure of the spacecraft. Communications equipment is placed on the despun section to provide for a stationary dual parabolic offset-feed array antenna system. The dual system is necessary to provide communications coverage during portions of maximum excursion on the halo orbit. Transmissions to the NASA Deep Space Network 34 m antenna include six channels (color video, two voice, scientific data from lunar payloads, satellite housekeeping and telemetry and uplinked commands) using the S- and X-bands. Four radioisotope thermoelectric generators (RTG's) provide a total of 1360 W to power onboard systems and any two of the four Hughes 13 cm ion thrusters at once. Output of the ion thrusters is approximately 17.8 mN each with xenon as the propellant. Presence of torques generated by solar pressure on the antenna dish require the addition of a 'skirt' extending from the spun section of the satellite for balance. Total mass of the satellite is approximately 900 kg at a cost of $130 million FY99.

Passive vs. Parachute System Architecture for Robotic Sample Return Vehicles

NASA Technical Reports Server (NTRS)

Maddock, Robert W.; Henning, Allen B.; Samareh, Jamshid A.

2016-01-01

The Multi-Mission Earth Entry Vehicle (MMEEV) is a flexible vehicle concept based on the Mars Sample Return (MSR) EEV design which can be used in the preliminary sample return mission study phase to parametrically investigate any trade space of interest to determine the best entry vehicle design approach for that particular mission concept. In addition to the trade space dimensions often considered (e.g. entry conditions, payload size and mass, vehicle size, etc.), the MMEEV trade space considers whether it might be more beneficial for the vehicle to utilize a parachute system during descent/landing or to be fully passive (i.e. not use a parachute). In order to evaluate this trade space dimension, a simplified parachute system model has been developed based on inputs such as vehicle size/mass, payload size/mass and landing requirements. This model works in conjunction with analytical approximations of a mission trade space dataset provided by the MMEEV System Analysis for Planetary EDL (M-SAPE) tool to help quantify the differences between an active (with parachute) and a passive (no parachute) vehicle concept.

JWST Mirror Technology Development

NASA Technical Reports Server (NTRS)

Stahl, H. Philip

2010-01-01

Since the initial Design Studies leading to JWST, Mirror Technology was identified as a (if not the) critical capability necessary to enable the next generation of large aperture space telescopes required to achieve the science goals of imaging the earliest galaxies and proto-galaxies after the big bang. Specific telescope architectures were explored via three independent design concept studies conducted during the summer of 1996. Achieving the desired science objectives required a never before demonstrated space telescope capability, one with an 8 meter class primary mirror that is diffraction limited at 2 micrometers and operating in deep space at temperatures well below 70K. Beryllium was identified in the NASA "Yardstick" design as the preferred material because of its ability to provide stable optical performance in the anticipated thermal environment as well as its excellent specific stiffness. Because of launch vehicle constraints, two very significant architectural constraints were placed upon the telescope: segmentation and areal density. Each of these directly resulted in specific technology capability requirements. First, because the maximum launch vehicle payload fairing diameter is approximately 4.5 meters, the only way to launch an 8 meter class mirror is to segment it, fold it and deploy it on orbit - resulting in actuation and control requirements. Second, because of launch vehicle mass limits, the primary mirror allocation was only 1000 kg - resulting in a maximum areal density specification of 20 kilograms per square meter.

The space shuttle payload planning working groups. Volume 5: Solar physics

NASA Technical Reports Server (NTRS)

1973-01-01

The findings of the Solar Physics working group of the space shuttle payload planning activity are presented. The areas to be investigated by the solar physics experiments are: (1) the production of mechanical energy in the subphotospheric layers and its transport and dissipation in the upper layers of the atmosphere, (2) the mass flux from the subphotospheric layers into the chromosphere and corona and beyond the solar wind, (3) solar activity and its relationship to magnetic fields, and (4) the production of solar flares. The approach to be followed in conducting the experiments and the equipment required are defined.

STS-39 OV-103 reaction control system (RCS) jets fire during onorbit maneuver

NASA Image and Video Library

1991-05-06

STS039-27-016 (28 April-6 May 1991) --- The Space Shuttle Discovery fires reaction control subsystem (RCS) thrusters in this 35mm frame, taken from inside the crew cabin. Seen in Discovery's payload bay are the tops of cannisters on the STP-1 payload, configured on the STS 39 Hitchhiker carrier; and the Air Force Program (AFP) 675 package. AFP-675 consists of the Cryogenic Infrared Radiance Instrumentation for Shuttle (CIRRIS)-1A; Far Ultraviolet Camera (FAR-UV) Experiment; Horizon Ultraviolet Program (HUP); Quadruple Ion Neutral Mass Spectrometer (QINMS); and the Uniformly Redundant Array (URA).

Green Propulsion Advancement and Infusion

NASA Technical Reports Server (NTRS)

Mulkey, Henry W.; Maynard, Andrew P.; Anflo, Kjell

2018-01-01

All space missions benefit from increased propulsion system performance allowing lower spacecraft launch mass, larger scientific payloads, or extended on-orbit lifetimes. Likewise, propellant candidates that offer significant reduction in personnel hazards and shorter payload processing present a more attractive propulsion subsystem solution. Aiming to reduce risk to potential infusion missions and fully comprehend the alternative propellant performance, the work presented herein represents many years of development and collaborative efforts to successfully align higher performance, low toxicity hydrazine alternatives into NASA Goddard Space Flight Center (GSFC) missions. The High Performance Green Propulsion (HPGP) technology is being considered for Science Mission Directorate (SMD) missions.

An advanced OBP-based payload operating in an asynchronous network for future data relay satellites utilising CCSDS-standard data structures

NASA Technical Reports Server (NTRS)

Grant, M.; Vernucci, A.

1991-01-01

A possible Data Relay Satellite System (DRSS) topology and network architecture is introduced. An asynchronous network concept, whereby each link (Inter-orbit, Inter-satellite, Feeder) is allowed to operate on its own clock, without causing loss of information, in conjunction with packet data structures, such as those specified by the CCSDS for advanced orbiting systems is discussed. A matching OBP payload architecture is described, highlighting the advantages provided by the OBP-based concept and then giving some indications on the OBP mass/power requirements.

Entry, Descent, and Landing Technology Concept Trade Study for Increasing Payload Mass to the Surface of Mars

NASA Technical Reports Server (NTRS)

Cruz, Juan R.; Cianciolo, Alicia D.; Powell, Richard W.; Simonsen, Lisa C.; Tolson, Robert H.

2005-01-01

A trade study was conducted that compared various entry, descent, and landing technologies and concepts for placing an 1,800 kg payload on the surface of Mars. The purpose of this trade study was to provide data, and make recommendations, that could be used in making decisions regarding which new technologies and concepts should be pursued. Five concepts were investigated, each using a different combination of new technologies: 1) a Baseline concept using the least new technologies, 2) Aerocapture and Entry from Orbit, 3) Inflatable Aeroshell, 4) Mid L/D Aeroshell-A (high ballistic coefficient), and 5) Mid L/D Aeroshell-B (low ballistic coefficient). All concepts were optimized to minimize entry mass subject to a common set of key requirements. These key requirements were: A) landing a payload mass of 1,800 kg, B) landing at an altitude 2.5 km above the MOLA areoid, C) landing with a descent rate of 2.5 m/s, and D) using a single launch vehicle available within the NASA Expendable Launch Vehicle Contract without resorting to in-space assembly. Additional constraints were implemented, some common to all concepts and others specific to the new technologies used. Among the findings of this study are the following observations. Concepts using blunt-body aeroshells (1, 2, and 3 above) had entry masses between 4,028 kg and 4,123 kg. Concepts using mid L/D aeroshells (4 and 5 above) were significantly heavier with entry masses of 5,292 kg (concept 4) and 4,812 kg (concept 5). This increased weight was mainly due to the aeroshell. Based on a comparison of the concepts it was recommended that: 1) re-qualified and/or improved TPS materials be developed, 2) large subsonic parachutes be qualified. Aerocapture was identified as a promising concept, but system issues beyond the scope of this study need to be investigated. Inflatable aeroshells were identified as a promising new technology, but they require additional technology maturation work. For the class of missions investigated in this trade study, mid L/D aeroshells were not competitive on an entry mass basis as compared to blunt-body aeroshells.

Electric Propulsion

NASA Astrophysics Data System (ADS)

Baggett, R.

2004-11-01

Next Generation Electric Propulsion (NGEP) technology development tasks are working towards advancing solar-powered electric propulsion systems and components to levels ready for transition to flight systems. Current tasks within NGEP include NASA's Evolutionary Xenon Thruster (NEXT), Carbon Based Ion Optics (CBIO), NSTAR Extended Life Test (ELT) and low-power Hall Effect thrusters. The growing number of solar electric propulsion options provides reduced cost and flexibility to capture a wide range of Solar System exploration missions. Benefits of electric propulsion systems over state-of-the-art chemical systems include increased launch windows, which reduce mission risk; increased deliverable payload mass for more science; and a reduction in launch vehicle size-- all of which increase the opportunities for New Frontiers and Discovery class missions. The Dawn Discovery mission makes use of electric propulsion for sequential rendezvous with two large asteroids (Vesta then Ceres), something not possible using chemical propulsion. NEXT components and thruster system under development have NSTAR heritage with significant increases in maximum power and Isp along with deep throttling capability to accommodate changes in input power over the mission trajectory. NEXT will produce engineering model system components that will be validated (through qualification-level and integrated system testing) and ready for transition to flight system development. NEXT offers Discovery, New Frontiers, Mars Exploration and outer-planet missions a larger deliverable payload mass and a smaller launch vehicle size. CBIO addresses the need to further extend ion thruster lifetime by using low erosion carbon-based materials. Testing of 30-cm Carbon-Carbon and Pyrolytic graphite grids using a lab model NSTAR thruster are complete. In addition, JPL completed a 1000 hr. life test on 30-cm Carbon-Carbon grids. The NSTAR ELT was a life time qualification test started in 1999 with a goal of 88 kg throughput of Xenon propellant. The test was intentionally terminated in 2003 after accumulating 233 kg throughput. The thruster has been completely disassembled and the conditions of all components documented. Because most of the NSTAR design features have been used in the NEXT thruster, the success of the ELT goes a long way toward qualifying NEXT by similarity Recent mission analyses for Discovery and New Frontiers class missions have also identified potential benefits of low-power, high thrust Hall Effect thrusters. Estimated to be ready for mission implementation by 2008, low-power Hall systems could increase mission capture for electric propulsion by greatly reducing propulsion cost, mass and complexity.

LAVA Subsystem Integration and Testing for the RESOLVE Payload of the Resource Prospector Mission: Mass Spectrometers and Gas Chromatography

NASA Technical Reports Server (NTRS)

Coan, Mary R.; Stewart, Elaine M.

2015-01-01

The Regolith and Environment Science & Oxygen and Lunar Volatile Extraction (RESOLVE) payload is part of Resource Prospector (RP) along with a rover and a lander that are expected to launch in 2020. RP will identify volatile elements that may be combined and collected to be used for fuel, air, and water in order to enable deeper space exploration. The Resource Prospector mission is a key part of In-Situ Resource Utilization (ISRU). The demand for this method of utilizing resources at the site of exploration is increasing due to the cost of resupply missions and deep space exploration goals. The RESOLVE payload includes the Lunar Advanced Volatile Analysis (LAVA) subsystem. The main instrument used to identify the volatiles evolved from the lunar regolith is the Gas Chromatograph-Mass Spectrometer (GC-MS). LAVA analyzes the volatiles emitted from the Oxygen and Volatile Extraction Node (OVEN) Subsystem. The objective of OVEN is to obtain, weigh, heat and transfer evolved gases to LAVA through the connection between the two subsystems called the LOVEN line. This paper highlights the work completed during a ten week internship that involved the integration, testing, data analysis, and procedure documentation of two candidate mass spectrometers for the LAVA subsystem in order to aid in determining which model to use for flight. Additionally, the examination of data from the integrated Resource Prospector '15 (RP' 15) field test will be presented in order to characterize the amount of water detected from water doped regolith samples.

A Reflight of the Explorer-1 Science Mission: The Montana EaRth Orbiting Pico Explorer (MEROPE)

NASA Astrophysics Data System (ADS)

Klumpar, D. M.; Obland, M.; Hunyadi, G.; Jepsen, S.; Larsen, B.; Kankelborg, C.; Hiscock, W.

2001-05-01

Montana State University's interdisciplinary Space Science and Engineering Laboratory (SSEL) under support from the Montana NASA Space Grant Consortium is engaged in an earth orbiting satellite student design and flight project. The Montana EaRth Orbiting Pico Explorer (MEROPE) will carry a modern-day reproduction of the scientific payload carried on Explorer-1. On February 1, 1958 the United States launched its first earth orbiting satellite carrying a 14 kg scientific experiment built by Professor James Van Allen's group at the State University of Iowa (now The University of Iowa). The MEROPE student satellite will carry a reproduction, using current-day technology, of the scientific payload flown on Explorer-1. The CubeSat-class satellite will use currently available, low cost technologies to produce a payload-carrying satellite with a total orbital mass of 1 kg in a volume of 1 cubic liter. The satellite is to be launched in late 2001 into a 600 km, 65° inclination orbit. MEROPE will utilize passive magnetic orientation for 2-axis attitude control. A central microprocessor provides timing, controls on-board operations and switching, and enables data storage. Body mounted GaAs solar arrays are expected to provide in excess of 1.5 W. to maintain battery charge and operate the bus and payload. The Geiger counter will be operated at approximately 50% duty cycle, primarily during transits of the earth's radiation belts. Data will be stored on board and transmitted approximately twice per day to a ground station located on the Bozeman campus of the Montana State University. Owing to the 65° inclination, the instrument will also detect the higher energy portion of the electron spectrum responsible for the production of the Aurora Borealis. This paper describes both the technical implementation and design of the satellite and its payload as well as the not inconsiderable task of large team organization and management. As of March 2001, the student team consists of four graduate students and approximately 45 undergraduates in fields including Physics, Engineering, Computer Sciences, Business, and Liberal Arts. Satellites of this class have the potential to lead to low-cost constellations of sciencecraft making coordinated measurements of the highly dynamic and spatially structured space environment. While key tradeoffs between resource needs and resource availability (e.g. power, telemetry, mass, volume, and cost) constrain payload sophistication, the tremendous advantages of having even simple dispersed multipoint measurements of the Geospace environment far outweigh the loss of payload sophistication in many instances.

Tanker Argus: Re-supply for a LEO Cryogenic Propellant Depot

NASA Astrophysics Data System (ADS)

St. Germain, B.; Olds, J.; Kokan, T.; Marcus, L.; Miller, J.

The Argus reusable launch vehicle (RLV) concept is a single-stage-to-orbit conical, winged bodied vehicle powered by two liquid hydrogen/liquid oxygen supercharged ejector ramjets. The 3rd generation Argus launch vehicle utilizes advanced vehicle technologies along with a Maglev launch assist track. A tanker version of the Argus RLV is envisioned to provide an economical means of providing liquid fuel and oxidizer to an orbiting low-Earth orbit (LEO) propellant depot. This depot could then provide propellant to various spacecraft, including reusable orbital transfer vehicles used to ferry space solar power satellites to geo-stationary orbit. Two different tanker Argus configurations were analyzed. The first simply places additional propellant tanks inside the payload bay of an existing Argus reusable launch vehicle. The second concept is a modified Argus RLV in which the payload bay is removed and the vehicle propellant tanks are stretched to hold extra propellant. An iterative conceptual design process was used to design both Argus vehicles. This process involves various disciplines including aerodynamics, trajectory analysis, weights &structures, propulsion, operations, safety, and cost/economics. The payload bay version of tanker Argus, which has a gross mass of 256.3MT, is designed to deliver a 9.07MT payload to LEO. This payload includes propellant and the tank structure required to secure this propellant in the payload bay. The modified, pure tanker version of Argus has a gross mass of 218.6MT and is sized to deliver a full 9.07MT of propellant to LEO. The economic analysis performed for this study involved the calculation of many factors including the design/development and recurring costs of each vehicle. These results were used along with other economic assumptions to determine the "per kilogram" cost of delivering propellant to orbit. The results show that for a given flight rate the "per kilogram" cost is cheaper for the pure tanker version of Argus. However, the main goal of this study was to determine at which flight rate would it be financially beneficial to spend more development money to modify an existing, sunk cost, payload bay version of Argus in order to create a more efficient pure tanker version. For flight rates greater than approximately 320 flights/year, there is only a small financial motivation to develop a pure tanker version. At this flight rate both versions of Argus are able to deliver propellant to LEO at an approximate cost of 375/kg.

Tether Transportation System Study

NASA Technical Reports Server (NTRS)

Bangham, M. E.; Lorenzini, E.; Vestal, L.

1998-01-01

The projected traffic to geostationary earth orbit (GEO) is expected to increase over the next few decades. At the same time, the cost of delivering payloads from the Earth's surface to low earth orbit (LEO) is projected to decrease, thanks in part to the Reusable Launch Vehicle (RLV). A comparable reduction in the cost of delivering payloads from LEO to GEO is sought. The use of in-space tethers, eliminating the requirement for traditional chemical upper stages and thereby reducing the launch mass, has been identified as such an alternative. Spinning tethers are excellent kinetic energy storage devices for providing the large delta vee's required for LEO to GEO transfer. A single-stage system for transferring payloads from LEO to GEO was proposed some years ago. The study results presented here contain the first detailed analyses of this proposal, its extension to a two-stage system, and the likely implementation of the operational system.

Delta capability for launch of communications satellites

NASA Technical Reports Server (NTRS)

Grimes, D. W.; Russell, W. A., Jr.; Kraft, J. D.

1982-01-01

The evolution of capabilities and the current performance levels of the Delta launch vehicle are outlined. The first payload was the Echo I passive communications satellite, weighing 179 lb, and placed in GEO in 1960. Emphasis since then has been to use off-the-shelf hardware where feasible. The latest version in the 3924 first stage, 3920 second stage, and Pam D apogee kick motor third stage. The Delta is presently equipped to place 2800 lb in GEO, as was proven with the 2717 lb Anik-D1 satellite. The GEO payload placement performance matches the Shuttle's, and work is therefore under way to enhance the Delta performance to handle more massive payloads. Installation of the Castor-IV solid motor separation system, thereby saving mass by utilizing compressed nitrogen, rather than mechanical thrusters to remove the strap-on boosters, is indicated, together with use of a higher performance propellant and a wider nose fairing.

KSC-08pd2650

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, technicians clean contamination from the Super Lightweight Interchangeable Carrier, or SLIC. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2651

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, a technician cleans contamination from the Super Lightweight Interchangeable Carrier, or SLIC. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2648

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, a technician cleans contamination from the Super Lightweight Interchangeable Carrier, or SLIC. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2649

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, a technician cleans contamination from the Super Lightweight Interchangeable Carrier, or SLIC. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2647

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, the Super Lightweight Interchangeable Carrier, or SLIC, is uncovered so that technicians can clean contaminants found earlier. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A will be removed. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

TOPEX satellite option study

NASA Technical Reports Server (NTRS)

1982-01-01

The basic design of the fleet satellite communication spacecraft (FLTSATCOM) can easily accommodate any of the three payload options for the ocean dynamic topography experiment (TOPEX). The principal mission requirements as well as the payload accommodations and communications systems needed for launching this payload are reviewed. The existing FLTSATCOM satellite design is identified and the approaches for the proposed propulsion system are described in addition to subsystems for mechanical; power; attitude and velocity control; and telemetry, tracking and control are described. The compatability of FLTSATCOM with the launch vehicle is examined and its capabilities vs TOPEX requirements are summarized. Undetermined changes needed to meet data storage, thermal control, and area to mass ratio requirements are discussed. Cost estimates are included for budgetary and planning purposes. The availability of the described design is assessed based on the continuing production of FLTSATCOM spacecraft during the schedule span planned for TOPEX.

Development of the Mars Observer Laser Altimeter (MOLA)

NASA Technical Reports Server (NTRS)

Johnson, Bertrand L., Jr

1993-01-01

The Mars Observer (MO) spacecraft payload scientific mission is to gather data on Martian global topography, gravity, weather, magnetic field and its interaction with the solar flux, surface chemistry, and mineralogy over one Mars year. In mid-1988 the need for a replacement altimeter as part of the payload complement arose. The Mars Observer Laser Altimeter (MOLA) was proposed by GSFC as an in-house effort and shortly afterward was 'conditionally' accepted. Constraints on funding, schedule, power, and mass were imposed with periodic reviews during the instrument development to authorize continuation. MOLA was designed, tested, and delivered in less than 36 months and integrated with the spacecraft. During spacecraft payload testing, the laser failed due to contamination in the laser cavity. In only 6 months, the laser was removed, rebuilt from spare parts, retested, and the instrument reassembled, realigned, requalified, and again delivered for spacecraft integration. Other aspects of the development of the MOLA are presented.

KSC-00pp1690

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- In the Space Station Processing Facility, the P6 integrated truss segment is lowered into the payload transport canister under the watchful eyes of the worker inside the canister as well as the workers on the sides. After being secured in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC00pp1689

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- Workers in the Space Station Processing Facility line up on the sides of the payload transport canister as an overhead crane moves the P6 integrated truss segment into position above it. After being placed in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1689

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- Workers in the Space Station Processing Facility line up on the sides of the payload transport canister as an overhead crane moves the P6 integrated truss segment into position above it. After being placed in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC00pp1690

NASA Image and Video Library

2000-11-10

KENNEDY SPACE CENTER, FLA. -- In the Space Station Processing Facility, the P6 integrated truss segment is lowered into the payload transport canister under the watchful eyes of the worker inside the canister as well as the workers on the sides. After being secured in the canister, the truss will be transported to Launch Pad 39B and the payload changeout room. Then it will be moved into Space Shuttle Endeavour’s payload bay for mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

STS-97 P6 truss moves to a payload transport canister

NASA Technical Reports Server (NTRS)

2000-01-01

As it travels across the Space Station Processing Facility, the P6 integrated truss segment passes over the two Italian-built Multi-Purpose Logistics Modules, Leonardo (right) and Raffaello (behind Leonardo). The P6 is being moved to a payload transport canister for transfer to Launch Pad 39B. There it will be placed in Endeavour'''s payload bay for launch on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station'''s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a '''blanket''' that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch is scheduled Nov. 30 at 10:06 p.m. EST.

Earth-Affecting Solar Causes Observatory (EASCO): Results of the Mission Concept Study

NASA Technical Reports Server (NTRS)

Gopalswamy, Natchimuthuk

2011-01-01

Coronal mass ejections (CMEs) corotating interaction regions (CIRs) are two large-scale structures that originate from the Sun and affect the heliosphere in general and Earth in particular. While CIRs are generally detected by in-situ plasma signatures, CMEs are remote-sensed when they are still close to the Sun. The current understanding of CMEs primarily come from the SOHO and STEREO missions. In spite of the enormous progress made, there are some serious deficiencies in these missions. For example, these missions did not carry all the necessary instruments (STEREO did not have a magnetograph; SOHO did not have in-situ magnetometer). From the Sun-Earth line, SOHO was not well-suited for observing Earth-directed CMEs because of the occulting disk. STEREO's angle with the Sun-Earth line is changing constantly, so only a limited number of Earth-directed CMEs were observed in profile. In order to overcome these difficulties, we proposed a news L5 mission concept known as the Earth-Affecting Solar Causes Observatory (EASCO). The mission concept was recently studied at the Mission Design Laboratory (MDL), NASA Goddard Space Flight Center. The aim of the MDL study was to see how the scientific payload consisting of ten instruments can be accommodated in the spacecraft bus, what propulsion system can transfer the payload to the Sun-Earth L5, and what launch vehicles are appropriate. The study found that all the ten instruments can be readily accommodated and can be launched using an intermediate size vehicle such as Taurus II with enhanced faring. The study also found that a hybrid propulsion system consisting of an ion thruster (using approximately 55 kg of Xenon) and hydrazine (approximately 10 kg) is adequate to place the payload at L5. The transfer will take about 2 years and the science mission will last for 4 years around the next solar maximum in 2025. The mission can be readily extended for another solar cycle to get a solar-cycle worth of data on Earth-affecting CMEs and CIRs. This paper provides a highlight of the MDL study results.

ISS External Payload Platform - a new opportunity for research in the space environment

NASA Astrophysics Data System (ADS)

Steimle, Christian; Pape, Uwe

The International Space Station (ISS) is a widely accepted platform for research activities in low Earth orbit. To a wide extent these activities are conducted in the pressurised laboratories of the station and less in the outside environment. Suitable locations outside the ISS are rare, existing facilities fully booked for the coming years. To overcome this limitation, an external payload platform accessible for small size payloads on a commercial basis will be launched to the ISS and installed on the Japanese Experiment Module External Facility (JEM-EF) in the third quarter of 2014 and will be ready to be used by the scientific community on a fully commercial basis. The new External Payload Platform (EPP) and its opportunities and constraints assessed regarding future research activities on-board the ISS. The small size platform is realised in a cooperation between the companies NanoRacks, Astrium North America in the United States, and Airbus Defence and Space in Germany. The hardware allows the fully robotic installation and operation of payloads. In the nominal mission scenario payload items are installed not later than one year after the signature of the contract, stay in operation for 15 weeks, and can be returned to the scientist thereafter. Payload items are transported among the pressurised cargo usually delivered to the station with various supply vehicles. Due to the high frequency of flights and the flexibility of the vehicle manifests the risk of a delay in the payload readiness can be mitigated by delaying to the next flight opportunity which on average is available not more than two months later. The mission is extra-ordinarily fast and of low cost in comparison to traditional research conducted on-board the ISS and can fit into short-term funding cycles available on national and multi-national levels. The size of the payload items is limited by handling constraints on-board the ISS. Therefore, the standard experiment payload size is a multiple of a 4U CubeSat, which demands miniaturised hardware solutions. But every payload can extensively use all ISS resources required: mass is not limited, power only limited by the payload heat radiation capability, the datalink is a USB 2.0 standard bus enabling a real-time and private data link. The new EPP transforms the station into a true laboratory in space with the capability to support research in various fields: exposure of biologic or material samples, experiments related to the radiation environment in low Earth orbit, and more.

Parametric Study of an Ablative TPS and Hot Structure Heatshield for a Mars Entry Capsule Vehicle

NASA Technical Reports Server (NTRS)

Langston, Sarah L.; Lang, Christapher G.; Samareh, Jamshid A.

2017-01-01

The National Aeronautics and Space Administration is planning to send humans to Mars. As part of the Evolvable Mars Campaign, different en- try vehicle configurations are being designed and considered for delivering larger payloads than have been previously sent to the surface of Mars. Mass and packing volume are driving factors in the vehicle design, and the thermal protection for planetary entry is an area in which advances in technology can offer potential mass and volume savings. The feasibility and potential benefits of a carbon-carbon hot structure concept for a Mars entry vehicle is explored in this paper. The windward heat shield of a capsule design is assessed for the hot structure concept as well as an ablative thermal protection system (TPS) attached to a honeycomb sandwich structure. Independent thermal and structural analyses are performed to determine the minimum mass design. The analyses are repeated for a range of design parameters, which include the trajectory, vehicle size, and payload. Polynomial response functions are created from the analysis results to study the capsule mass with respect to the design parameters. Results from the polynomial response functions created from the thermal and structural analyses indicate that the mass of the capsule was higher for the hot structure concept as compared to the ablative TPS for the parameter space considered in this study.

A UV LED-based Charge Management System for LISA

NASA Astrophysics Data System (ADS)

Conklin, John W.; Chilton, Andrew; Olatunde, Taiwo Janet; Apple, Stephen; Parry, Samantha; Ciani, Giacomo; Wass, Peter; Mueller, Guido

2018-01-01

The Laser Interferometer Space Antenna (LISA) will be the first space instrument to observe gravitational waves in the millihertz frequency band. LISA consists of three Sun-orbiting spacecraft that form an equilateral triangle, with each side measuring 2.5 million kilometers in length. Each spacecraft houses two free-floating test masses, which are protected from all disturbing forces so that they follow pure geodesics in spacetime. A drag-free control system commands micronewton thrusters to force the spacecraft to fly in formation with the test masses and laser interferometers measure the minute variations in the distance, or light travel time, between these free-falling test masses caused by gravitational waves. The LISA observatory, with a planned launch in the early 2030s, is led by the European Space Agency with significant contributions from NASA. Recently, NASA has initiated strategic investments in key LISA technologies that will likely become U.S. flight hardware contributions to this ground-breaking mission. One of these payload elements is the Charge Management System (CMS), which controls the electric potential of the test masses relative to their housings to reduce spurious force noise acting on the test masses to below the required level. This talk, presented by University of Florida team that leads the CMS development, will describe this vital U.S. contribution to the LISA mission in the context of the envisioned LISA payload architecture and its in-flight sensitivity to gravitational waves.

77 FR 21619 - Office of Commercial Space Transportation; Notice of Intent To Prepare an Environmental Impact...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-04-10

... rocket propellant-1 or refined petroleum-1 (RP-1), as propellants to carry payloads into orbit. The... tank with a maximum propellant (RP-1 and LOX) load of approximately 6,900 gallons. As part of the... processing-hangar, a launch pad and stand with its associated flame duct, propellant storage and handling...

Gaia challenging performances verification: combination of spacecraft models and test results

NASA Astrophysics Data System (ADS)

Ecale, Eric; Faye, Frédéric; Chassat, François

2016-08-01

To achieve the ambitious scientific objectives of the Gaia mission, extremely stringent performance requirements have been given to the spacecraft contractor (Airbus Defence and Space). For a set of those key-performance requirements (e.g. end-of-mission parallax, maximum detectable magnitude, maximum sky density or attitude control system stability), this paper describes how they are engineered during the whole spacecraft development process, with a focus on the end-to-end performance verification. As far as possible, performances are usually verified by end-to-end tests onground (i.e. before launch). However, the challenging Gaia requirements are not verifiable by such a strategy, principally because no test facility exists to reproduce the expected flight conditions. The Gaia performance verification strategy is therefore based on a mix between analyses (based on spacecraft models) and tests (used to directly feed the models or to correlate them). Emphasis is placed on how to maximize the test contribution to performance verification while keeping the test feasible within an affordable effort. In particular, the paper highlights the contribution of the Gaia Payload Module Thermal Vacuum test to the performance verification before launch. Eventually, an overview of the in-flight payload calibration and in-flight performance verification is provided.

Dynamic analysis of astronaut motions in microgravity: Applications for Extravehicular Activity (EVA)

NASA Technical Reports Server (NTRS)

Newman, Dava J.

1995-01-01

Simulations of astronaut motions during extravehicular activity (EVA) tasks were performed using computational multibody dynamics methods. The application of computational dynamic simulation to EVA was prompted by the realization that physical microgravity simulators have inherent limitations: viscosity in neutral buoyancy tanks; friction in air bearing floors; short duration for parabolic aircraft; and inertia and friction in suspension mechanisms. These limitations can mask critical dynamic effects that later cause problems during actual EVA's performed in space. Methods of formulating dynamic equations of motion for multibody systems are discussed with emphasis on Kane's method, which forms the basis of the simulations presented herein. Formulation of the equations of motion for a two degree of freedom arm is presented as an explicit example. The four basic steps in creating the computational simulations were: system description, in which the geometry, mass properties, and interconnection of system bodies are input to the computer; equation formulation based on the system description; inverse kinematics, in which the angles, velocities, and accelerations of joints are calculated for prescribed motion of the endpoint (hand) of the arm; and inverse dynamics, in which joint torques are calculated for a prescribed motion. A graphical animation and data plotting program, EVADS (EVA Dynamics Simulation), was developed and used to analyze the results of the simulations that were performed on a Silicon Graphics Indigo2 computer. EVA tasks involving manipulation of the Spartan 204 free flying astronomy payload, as performed during Space Shuttle mission STS-63 (February 1995), served as the subject for two dynamic simulations. An EVA crewmember was modeled as a seven segment system with an eighth segment representing the massive payload attached to the hand. For both simulations, the initial configuration of the lower body (trunk, upper leg, and lower leg) was a neutral microgravity posture. In the first simulation, the payload was manipulated around a circular trajectory of 0.15 m radius in 10 seconds. It was found that the wrist joint theoretically exceeded its ulnal deviation limit by as much as 49. 8 deg and was required to exert torques as high as 26 N-m to accomplish the task, well in excess of the wrist physiological limit of 12 N-m. The largest torque in the first simulation, 52 N-m, occurred in the ankle joint. To avoid these problems, the second simulation placed the arm in a more comfortable initial position and the radius and speed of the circular trajectory were reduced by half. As a result, the joint angles and torques were reduced to values well within their physiological limits. In particular, the maximum wrist torque for the second simulation was only 3 N-m and the maximum ankle torque was only 6 N-m.

Rockot - a new cost effective launcher for small satellites

NASA Astrophysics Data System (ADS)

Mosenkis, Regina

1996-01-01

Daimler-Benz Aerospace of Germany and the Russian Khrunichev State Research and Production Space Center have formed a jointly owned EUROCKOT Launch Services GmbH to offer worldwide cost effective launch services for the ROCKOT launch vehicle. ROCKOT, produced by Khrunichev, builder of the famous PROTON launcher, aims at the market of small and medium size satellites ranging from 300 to 1800 kg to be launched into low earth or sunsynchronous orbits. These comprize scientific, earth observation and polar meteorological satellites as well as the new generation of small communication satellites in low earth orbits, known as the ``Constellations''. ROCKOT is a three stage liquid propellant launch vehicle, composed of a former Russian SS 19 strategic missile, which has been withdrawn from military use, and a highly sophisticated, flight-proven upper stage named Breeze, which is particularly suited for a variety of civic and commercial space applications. Usable payload envelope has a length of 4.75 meters and a maximum diameter of 2.26 meters for accomodating the payload within the payload fairing. ROCKOT can also accomodate multiple payloads which can be deployed into the same or different orbits. So far ROCKOT has been successfully launched three times from Baikonur. The commercial launch services on ROCKOT from the Plesetsk launch site, Russia, will begin in 1997 and will be available worldwide at a highly competitive price.

An improved input shaping design for an efficient sway control of a nonlinear 3D overhead crane with friction

NASA Astrophysics Data System (ADS)

Maghsoudi, Mohammad Javad; Mohamed, Z.; Sudin, S.; Buyamin, S.; Jaafar, H. I.; Ahmad, S. M.

2017-08-01

This paper proposes an improved input shaping scheme for an efficient sway control of a nonlinear three dimensional (3D) overhead crane with friction using the particle swarm optimization (PSO) algorithm. Using this approach, a higher payload sway reduction is obtained as the input shaper is designed based on a complete nonlinear model, as compared to the analytical-based input shaping scheme derived using a linear second order model. Zero Vibration (ZV) and Distributed Zero Vibration (DZV) shapers are designed using both analytical and PSO approaches for sway control of rail and trolley movements. To test the effectiveness of the proposed approach, MATLAB simulations and experiments on a laboratory 3D overhead crane are performed under various conditions involving different cable lengths and sway frequencies. Their performances are studied based on a maximum residual of payload sway and Integrated Absolute Error (IAE) values which indicate total payload sway of the crane. With experiments, the superiority of the proposed approach over the analytical-based is shown by 30-50% reductions of the IAE values for rail and trolley movements, for both ZV and DZV shapers. In addition, simulations results show higher sway reductions with the proposed approach. It is revealed that the proposed PSO-based input shaping design provides higher payload sway reductions of a 3D overhead crane with friction as compared to the commonly designed input shapers.

A Vibration Isolation System for Use in a Large Thermal Vacuum Test Facility

NASA Technical Reports Server (NTRS)

Hershfeld, Donald; VanCampen, Julie

2002-01-01

A thermal vacuum payload platform that is isolated from background vibration is required to support the development of future instruments for Hubble Space Telescope (HST) and the Next Generation Space Telescope (NGST) at the Goddard Space Flight Center (GSFC). Because of the size and weight of the thermal/vacuum facility in which the instruments are tested, it is not practical to isolate the entire facility externally. Therefore, a vibration isolation system has been designed and fabricated to be installed inside the chamber. The isolation system provides a payload interface of 3.05 m (10 feet) in diameter and is capable of supporting a maximum payload weight of 4536 kg (10,000 Lbs). A counterweight system has been included to insure stability of payloads having high centers of gravity. The vibration isolation system poses a potential problem in that leakage into the chamber could compromise the ability to maintain vacuum. Strict specifications were imposed on the isolation system design to minimize leakage. Vibration measurements, obtained inside the chamber, prior to installing the vibration isolation system, indicated levels in all axes of approximately 1 milli-g at about 20 Hz. The vibration isolation system was designed to provide a minimum attenuation of 40 dB to these levels. This paper describes the design and testing of this unique vibration isolation system. Problems with leakage and corrective methods are presented. Isolation performance results are also presented.

Inficon Transpector MPH Mass Spectrometer Random Vibration Test Report

NASA Technical Reports Server (NTRS)

Santiago-Bond, Jo; Captain, Janine

2015-01-01

The purpose of this test report is to summarize results from the vibration testing of the INFICON Transpector MPH100M model Mass Spectrometer. It also identifies requirements satisfied, and procedures used in the test. As a payload of Resource Prospector, it is necessary to determine the survivability of the mass spectrometer to proto-qualification level random vibration. Changes in sensitivity of the mass spectrometer can be interpreted as a change in alignment of the instrument. The results of this test will be used to determine any necessary design changes as the team moves forward with flight design.

A survey of propulsion options for cargo and piloted missions to Mars.

PubMed

Sankaran, K; Cassady, L; Kodys, A D; Choueiri, E Y

2004-05-01

In this paper, high-power electric propulsion options are surveyed in the context of cargo and piloted missions to Mars. A low-thrust trajectory optimization program (raptor) is utilized to analyze this mission. Candidate thrusters are chosen based upon demonstrated performance in the laboratory. Hall, self-field magnetoplasmadynamic (MPDT), self-field lithium Lorentz force accelerator (LiLFA), arcjet, and applied-field LiLFA systems are considered for this mission. In this first phase of the study, all thrusters are assumed to operate at a single power level (regardless of the efficiency-power curve), and the thruster specific mass and power plant specific mass are taken to be the same for all systems. Under these assumptions, for a 7.5 MW, 60 mT payload, piloted mission, the self-field LiLFA results in the shortest trip time (340 days) with a reasonable propellant mass fraction of 57% (129 mT). For a 150 kW, 9 mT payload, cargo mission, both the applied-field LiLFA and the Hall thruster seem reasonable choices with propellant mass fractions of 42 to 45%(7 to 8 mT). The Hall thrusters provide better trip times (530-570 days) compared to the applied-field LiLFA (710 days) for the relatively less demanding mission.

Integrated Payload Data Handling Systems Using Software Partitioning

NASA Astrophysics Data System (ADS)

Taylor, Alun; Hann, Mark; Wishart, Alex

2015-09-01

An integrated Payload Data Handling System (I-PDHS) is one in which multiple instruments share a central payload processor for their on-board data processing tasks. This offers a number of advantages over the conventional decentralised architecture. Savings in payload mass and power can be realised because the total processing resource is matched to the requirements, as opposed to the decentralised architecture here the processing resource is in effect the sum of all the applications. Overall development cost can be reduced using a common processor. At individual instrument level the potential benefits include a standardised application development environment, and the opportunity to run the instrument data handling application on a fully redundant and more powerful processing platform [1]. This paper describes a joint program by SCISYS UK Limited, Airbus Defence and Space, Imperial College London and RAL Space to implement a realistic demonstration of an I-PDHS using engineering models of flight instruments (a magnetometer and camera) and a laboratory demonstrator of a central payload processor which is functionally representative of a flight design. The objective is to raise the Technology Readiness Level of the centralised data processing technique by address the key areas of task partitioning to prevent fault propagation and the use of a common development process for the instrument applications. The project is supported by a UK Space Agency grant awarded under the National Space Technology Program SpaceCITI scheme. [1].

LEO to GEO (and Beyond) Transfers Using High Power Solar Electric Propulsion (HP-SEP)

NASA Technical Reports Server (NTRS)

Loghry, Christopher S.; Oleson, Steven R.; Woytach, Jeffrey M.; Martini, Michael C.; Smith, David A.; Fittje, James E.; Gyekenyesi, John Z.; Colozza, Anthony J.; Fincannon, James; Bogner, Aimee;

Geostationary payload concepts for personal satellite communications

NASA Technical Reports Server (NTRS)

Benedicto, J.; Rinous, P.; Roberts, I.; Roederer, A.; Stojkovic, I.

1993-01-01

This paper reviews candidate satellite payload architectures for systems providing world-wide communication services to mobile users equipped with hand-held terminals based on large geostationary satellites. There are a number of problems related to the payload architecture, on-board routing and beamforming, and the design of the S-band Tx and L-band Rx antenna and front ends. A number of solutions are outlined, based on trade-offs with respect to the most significant performance parameters such as capacity, G/T, flexibility of routing traffic to beams and re-configuration of the spot-beam coverage, and payload mass and power. Candidate antenna and front-end configurations were studied, in particular direct radiating arrays, arrays magnified by a reflector and active focused reflectors with overlapping feed clusters for both transmit (multimax) and receive (beam synthesis). Regarding the on-board routing and beamforming sub-systems, analog techniques based on banks of SAW filters, FET or CMOS switches and cross-bar fixed and variable beamforming are compared with a hybrid analog/digital approach based on Chirp Fourier Transform (CFT) demultiplexer combined with digital beamforming or a fully digital processor implementation, also based on CFT demultiplexing.

High performance photonic ADC for space applications

NASA Astrophysics Data System (ADS)

Pantoja, S.; Piqueras, M. A.; Villalba, P.; Martínez, B.; Rico, E.

2017-11-01

The flexibility required for future telecom payloads will require of more digital processing capabilities, moving from conventional analogue repeaters to more advanced and efficient analog subsystems or DSPbased solutions. Aggregate data throughputs will have to be handled onboard, creating the need for effective, ADC/DSP and DSP/DAC high speed links. Broadband payloads will have to receive, route and retransmit hundreds of channels and need to be designed so as to meet such requirements of larger bandwidth, system transparency and flexibility.[1][2] One important device in these new architectures is analog to digital converter (ADC) and its equivalent digital to analog converter (DAC). These will be the in/out interface for the use of digital processing in order to provide flexible beam to beam connectivity and variable bandwidth allocation. For telecom payloads having a large number of feeds and thus a large number of converters the mass and consumption of the mixer stage has become significant. Moreover, the inclusion of ADCs in the payload presents new trade-offs in design (jitter, quantization noise, ambiguity). This paper deals with an alternative solution of these two main problems with the exploitation of photonic techniques.

Human Mars Entry, Descent and Landing Architectures Study Overview

NASA Technical Reports Server (NTRS)

Polsgrove, Tara T.; Dwyer Cianciolo, Alicia

2016-01-01

Landing humans on Mars will require entry, descent and landing (EDL) capability beyond the current state of the art. Nearly twenty times more delivered payload and an order of magnitude improvement in precision landing capability will be necessary. Several EDL technologies capable of meeting the human class payload delivery requirements are being considered. The EDL technologies considered include low lift-to-drag vehicles like Hypersonic Inflatable Aerodynamic Decelerators (HIAD), Adaptable Deployable Entry and Placement Technology (ADEPT), and mid range lift-to-drag vehicles like rigid aeroshell configurations. To better assess EDL technology options and sensitivities to future human mission design variations, a series of design studies has been conducted. The design studies incorporate EDL technologies with conceptual payload arrangements defined by the Evolvable Mars Campaign to evaluate the integrated system with higher fidelity than have been performed to date. This paper describes the results of the design studies for a lander design using the HIAD, ADEPT and rigid shell entry technologies and includes system and subsystem design details including mass and power estimates. This paper will review the point design for three entry configurations capable of delivering a 20 t human class payload to the surface of Mars.

Target selection and mass estimation for manned NEO exploration using a baseline mission design

NASA Astrophysics Data System (ADS)

Boden, Ralf C.; Hein, Andreas M.; Kawaguchi, Junichiro

2015-06-01

In recent years Near-Earth Objects (NEOs) have received an increased amount of interest as a target for human exploration. NEOs offer scientifically interesting targets, and at the same time function as a stepping stone for achieving future Mars missions. The aim of this research is to identify promising targets from the large number of known NEOs that qualify for a manned sample-return mission with a maximum duration of one year. By developing a baseline mission design and a mass estimation model, mission opportunities are evaluated based on on-orbit mass requirements, safety considerations, and the properties of the potential targets. A selection of promising NEOs is presented and the effects of mission requirements and restrictions are discussed. Regarding safety aspects, the use of free-return trajectories provides the lowest on-orbit mass, when compared to an alternative design that uses system redundancies to ensure return of the spacecraft to Earth. It is discovered that, although a number of targets are accessible within the analysed time frame, no NEO offers both easy access and high incentive for its exploration. Under the discussed aspects a first human exploration mission going beyond the vicinity of Earth will require a trade off between targets that provide easy access and those that are of scientific interest. This lack of optimal mission opportunities can be seen in the small number of only 4 NEOs that meet all requirements for a sample-return mission and remain below an on-orbit mass of 500 metric Tons (mT). All of them require a mass between 315 and 492 mT. Even less ideal, smaller asteroids that are better accessible require an on-orbit mass that exceeds the launch capability of future heavy lift vehicles (HLV) such as SLS by at least 30 mT. These mass requirements show that additional efforts are necessary to increase the number of available targets and reduce on-orbit mass requirements through advanced mission architectures. The need for on-orbit assembly also becomes apparent, as availability of a HLV alone does not provide sufficient payload capabilities for any manned mission targeting NEOs.

Data and performances of selected aircraft and rotorcraft

NASA Astrophysics Data System (ADS)

Filippone, Antonio

2000-11-01

The purpose of this article is to provide a synthetic and comparative view of selected aircraft and rotorcraft (nearly 300 of them) from past and present. We report geometric characteristics of wings (wing span, areas, aspect-ratios, sweep angles, dihedral/anhedral angles, thickness ratios at root and tips, taper ratios) and rotor blades (type of rotor, diameter, number of blades, solidity, rpm, tip Mach numbers); aerodynamic data (drag coefficients at zero lift, cruise and maximum absolute glide ratio); performances (wing and disk loadings, maximum absolute Mach number, cruise Mach number, service ceiling, rate of climb, centrifugal acceleration limits, maximum take-off weight, maximum payload, thrust-to-weight ratios). There are additional data on wing types, high-lift devices, noise levels at take-off and landing. The data are presented on tables for each aircraft class. A graphic analysis offers a comparative look at all types of data. Accuracy levels are provided wherever available.

Design and evaluation of a payload to support plant growth onboard COMET 1

NASA Technical Reports Server (NTRS)

Hoehn, A.; Kliss, M. H.; Luttges, M. W.; Robinson, M. C.; Stodieck, L. S.

1992-01-01

The paper describes the design and the operation principles of the Plant Module for Autonomous Space Support (P-MASS), designed to provide life support for a variety of plants, algae, and bacteria in low earth orbit during the maiden flight of COMET-1, scheduled for 1993. During flight (scheduled to continue for 30 days), both color video images and collected environmental data (including light intensity, temperature, relative humidity, CO2 and O2 concentrations, soil moisture, and nutrients released) will be downlinked to earth several times a day. These data will also be stored within the payload and retrieved from it after reentry and recovery.

Primary propulsion of electrothermal, ion, and chemical systems for space-based radar orbit transfer

NASA Technical Reports Server (NTRS)

Wang, S.-Y.; Staiger, P. J.

1985-01-01

An orbit transfer mission concept has been studied for a Space-Based Radar (SBR) where 40 kW required for radar operation is assumed available for orbit transfer propulsion. Arcjet, pulsed electrothermal (PET), ion, and storable chemical systems are considered for the primary propulsion. Transferring two SBR per shuttle flight to 1112 km/60 deg using eiectrical propulsion systems offers an increased payload at the expense of increased trip time, up to 2000 kg each, which may be critical for survivability. Trade offs between payload mass, transfer time, launch site, inclination, and height of parking orbits are presented.

An adaptive controller for enhancing operator performance during teleoperation

NASA Technical Reports Server (NTRS)

Carignan, Craig R.; Tarrant, Janice M.; Mosier, Gary E.

1989-01-01

An adaptive controller is developed for adjusting robot arm parameters while manipulating payloads of unknown mass and inertia. The controller is tested experimentally in a master/slave configuration where the adaptive slave arm is commanded via human operator inputs from a master. Kinematically similar six-joint master and slave arms are used with the last three joints locked for simplification. After a brief initial adaptation period for the unloaded arm, the slave arm retrieves different size payloads and maneuvers them about the workspace. Comparisons are then drawn with similar tasks where the adaptation is turned off. Several simplifications of the controller dynamics are also addressed and experimentally verified.

A study of cooling time reduction of interferometric cryogenic gravitational wave detectors using a high-emissivity coating

DOE Office of Scientific and Technical Information (OSTI.GOV)

Sakakibara, Y.; Yamamoto, K.; Chen, D.

In interferometric cryogenic gravitational wave detectors, there are plans to cool mirrors and their suspension systems (payloads) in order to reduce thermal noise, that is, one of the fundamental noise sources. Because of the large payload masses (several hundred kg in total) and their thermal isolation, a cooling time of several months is required. Our calculation shows that a high-emissivity coating (e.g. a diamond-like carbon (DLC) coating) can reduce the cooling time effectively by enhancing radiation heat transfer. Here, we have experimentally verified the effect of the DLC coating on the reduction of the cooling time.

Primary propulsion of electrothermal, ion and chemical systems for space-based radar orbit transfer

NASA Technical Reports Server (NTRS)

Wang, S. Y.; Staiger, P. J.

1985-01-01

An orbit transfer mission concept has been studied for a Space-Based Radar (SBR) where 40 kW required for radar operation is assumed available for orbit transfer propulsion. Arcjet, pulsed electrothermal (PET), ion, and storable chemical systems are considered for the primary propulsion. Transferring two SBR per shuttle flight to 1112 km/60 deg using electrical propulsion systems offers an increased payload at the expense of increased trip time, up to 2000 kg each, which may be critical for survivability. Trade offs between payload mass, transfer time, launch site, inclination, and height of parking orbits are presented.

High Altitude Flight Test of a 40-Foot Diameter (12.2 meter) Ringsail Parachute at Deployment Mach Number of 2.95

NASA Technical Reports Server (NTRS)

Eckstrom, Clinton V.

1970-01-01

A 40-foot-nominal-diameter (12.2-meter) modified ringsail parachute was flight tested as part of the NASA Supersonic High Altitude Parachute Experiment (SHAPE) program. The 41-pound (18.6-kg) test parachute system was deployed from a 239.5-pound (108.6-kg) instrumented payload by means of a deployment mortar when the payload was at an altitude of 171,400 feet (52.3 km), a Mach number of 2.95, and a free-stream dynamic pressure of 9.2 lb/sq ft (440 N/m(exp 2)). The parachute deployed properly, suspension line stretch occurring 0.54 second after mortar firing with a resulting snatch-force loading of 932 pounds (4146 newtons). The maximum loading due to parachute opening was 5162 pounds (22 962 newtons) at 1.29 seconds after mortar firing. The first near full inflation of the canopy at 1.25 seconds after mortar firing was followed immediately by a partial collapse and subsequent oscillations of frontal area until the system had decelerated to a Mach number of about 1.5. The parachute then attained a shape that provided full drag area. During the supersonic part of the test, the average axial-force coefficient varied from a minimum of about 0.24 at a Mach number of 2.7 to a maximum of 0.54 at a Mach number of 1.1. During descent under subsonic conditions, the average effective drag coefficient was 0.62 and parachute-payload oscillation angles averaged about &loo with excursions to +/-20 degrees. The recovered parachute was found to have slight damage in the vent area caused by the attached deployment bag and mortar lid.

A concept for NASA's Mars 2016 astrobiology field laboratory.

PubMed

Beegle, Luther W; Wilson, Michael G; Abilleira, Fernando; Jordan, James F; Wilson, Gregory R

2007-08-01

The Mars Program Plan includes an integrated and coordinated set of future candidate missions and investigations that meet fundamental science objectives of NASA and the Mars Exploration Program (MEP). At the time this paper was written, these possible future missions are planned in a manner consistent with a projected budget profile for the Mars Program in the next decade (2007-2016). As with all future missions, the funding profile depends on a number of factors that include the exact cost of each mission as well as potential changes to the overall NASA budget. In the current version of the Mars Program Plan, the Astrobiology Field Laboratory (AFL) exists as a candidate project to determine whether there were (or are) habitable zones and life, and how the development of these zones may be related to the overall evolution of the planet. The AFL concept is a surface exploration mission equipped with a major in situ laboratory capable of making significant advancements toward the Mars Program's life-related scientific goals and the overarching Vision for Space Exploration. We have developed several concepts for the AFL that fit within known budget and engineering constraints projected for the 2016 and 2018 Mars mission launch opportunities. The AFL mission architecture proposed here assumes maximum heritage from the 2009 Mars Science Laboratory (MSL). Candidate payload elements for this concept were identified from a set of recommendations put forth by the Astrobiology Field Laboratory Science Steering Group (AFL SSG) in 2004, for the express purpose of identifying overall rover mass and power requirements for such a mission. The conceptual payload includes a Precision Sample Handling and Processing System that would replace and augment the functionality and capabilities provided by the Sample Acquisition Sample Processing and Handling system that is currently part of the 2009 MSL platform.

Advanced propulsion for LEO and GEO platforms

NASA Technical Reports Server (NTRS)

Sovey, James S.; Pidgeon, David J.

1990-01-01

Mission requirements and mass savings applicable to specific low earth orbit and geostationary earth orbit platforms using three highly developed propulsion systems are described. Advanced hypergolic bipropellant thrusters and hydrazine arcjets can provide about 11 percent additional instrument payload to 14,000 kg LEO platforms. By using electric propulsion on a 8,000 kg class GEO platform, mass savings in excess of 15 percent of the beginning-of-life platform mass are obtained. Effects of large, advanced technology solar arrays and antennas on platform propulsion requirements are also discussed.

Two-step Laser Time-of-Flight Mass Spectrometry to Elucidate Organic Diversity in Planetary Surface Materials.

NASA Technical Reports Server (NTRS)

Getty, Stephanie A.; Brinckerhoff, William B.; Cornish, Timothy; Li, Xiang; Floyd, Melissa; Arevalo, Ricardo Jr.; Cook, Jamie Elsila; Callahan, Michael P.

2013-01-01

Laser desorption/ionization time-of-flight mass spectrometry (LD-TOF-MS) holds promise to be a low-mass, compact in situ analytical capability for future landed missions to planetary surfaces. The ability to analyze a solid sample for both mineralogical and preserved organic content with laser ionization could be compelling as part of a scientific mission pay-load that must be prepared for unanticipated discoveries. Targeted missions for this instrument capability include Mars, Europa, Enceladus, and small icy bodies, such as asteroids and comets.

Preferential heating of light ions during an ionospheric Ar(+) injection experiment

NASA Technical Reports Server (NTRS)

Pollock, C. J.; Chandler, M. O.; Moore, T. E.; Arnoldy, R. L.; Kintner, P. M.; Chesney, S.; Cahill, L. J., Jr.

1995-01-01

The Argon Release for Controlled Studies (ARCS) 4 sounding rocket was launched northward into high altitude from Poker Flat Research Range on February 23, 1990. The vehicle crossed geomagnetic field lines containing discrete auroral activity. An instrumented subpayload released 100-eV and 200-eV Ar(+) ion beams sequentially, in a direction largely perpendicular to both the local geomagnetic field and the subpayload spin axis. The instrumented main payload was separated along field lines from the beam emitting subpayload by a distance which increased at a steady rate of approximately 2.4 m/s. Three dimensional mass spectrometric ion observations of ambient H(+) and O(+) ions, obtained on board the main payload, are presented. Main payload electric field observations in the frequency range 0-16 kHz, are also presented. These observations are presented to demonstrate the operation of transverse ion acceleration, which was differential with respect to ion mass, primarily during 100-eV beam operations. The preferential transverse acceleration of ambient H(+) ions, as compared with ambient O(+) ions, during the second, third, fourth, and fifth 100-eV beam operations, is attributed to a resonance among the injected Ar(+) ions, beam-generated lower hybrid waves, and H(+) ions in the tail of the ambient thermal distribution. This work provides experimental support of processes predicted by previously published theory and simulations.

Prompt merger collapse and the maximum mass of neutron stars.

PubMed

Bauswein, A; Baumgarte, T W; Janka, H-T

2013-09-27

We perform hydrodynamical simulations of neutron-star mergers for a large sample of temperature-dependent nuclear equations of state and determine the threshold mass above which the merger remnant promptly collapses to form a black hole. We find that, depending on the equation of state, the threshold mass is larger than the maximum mass of a nonrotating star in isolation by between 30 and 70 percent. Our simulations also show that the ratio between the threshold mass and maximum mass is tightly correlated with the compactness of the nonrotating maximum-mass configuration. We speculate on how this relation can be used to derive constraints on neutron-star properties from future observations.

SAM-CAAM: A Concept for Acquiring Systematic Aircraft Measurements to Characterize Aerosol Air Masses

DOE Office of Scientific and Technical Information (OSTI.GOV)

Kahn, Ralph A.; Berkoff, Tim A.; Brock, Charles

A modest operational program of systematic aircraft measurements can resolve key satellite aerosol data record limitations. Satellite observations provide frequent global aerosol amount maps but offer only loose aerosol property constraints needed for climate and air quality applications. In this paper, we define and illustrate the feasibility of flying an aircraft payload to measure key aerosol optical, microphysical, and chemical properties in situ. The flight program could characterize major aerosol airmass types statistically, at a level of detail unobtainable from space. It would 1) enhance satellite aerosol retrieval products with better climatology assumptions and 2) improve translation between satellite-retrieved opticalmore » properties and species-specific aerosol mass and size simulated in climate models to assess aerosol forcing, its anthropogenic components, and other environmental impacts. As such, Systematic Aircraft Measurements to Characterize Aerosol Air Masses (SAM-CAAM) could add value to data records representing several decades of aerosol observations from space; improve aerosol constraints on climate modeling; help interrelate remote sensing, in situ, and modeling aerosol-type definitions; and contribute to future satellite aerosol missions. Fifteen required variables are identified and four payload options of increasing ambition are defined to constrain these quantities. “Option C” could meet all the SAM-CAAM objectives with about 20 instruments, most of which have flown before, but never routinely several times per week, and never as a group. Aircraft integration and approaches to data handling, payload support, and logistical considerations for a long-term, operational mission are discussed. Finally, SAM-CAAM is feasible because, for most aerosol sources and specified seasons, particle properties tend to be repeatable, even if aerosol loading varies.« less

SAM-CAAM: A Concept for Acquiring Systematic Aircraft Measurements to Characterize Aerosol Air Masses

DOE PAGES

Kahn, Ralph A.; Berkoff, Tim A.; Brock, Charles; ...

2017-10-30

A modest operational program of systematic aircraft measurements can resolve key satellite aerosol data record limitations. Satellite observations provide frequent global aerosol amount maps but offer only loose aerosol property constraints needed for climate and air quality applications. In this paper, we define and illustrate the feasibility of flying an aircraft payload to measure key aerosol optical, microphysical, and chemical properties in situ. The flight program could characterize major aerosol airmass types statistically, at a level of detail unobtainable from space. It would 1) enhance satellite aerosol retrieval products with better climatology assumptions and 2) improve translation between satellite-retrieved opticalmore » properties and species-specific aerosol mass and size simulated in climate models to assess aerosol forcing, its anthropogenic components, and other environmental impacts. As such, Systematic Aircraft Measurements to Characterize Aerosol Air Masses (SAM-CAAM) could add value to data records representing several decades of aerosol observations from space; improve aerosol constraints on climate modeling; help interrelate remote sensing, in situ, and modeling aerosol-type definitions; and contribute to future satellite aerosol missions. Fifteen required variables are identified and four payload options of increasing ambition are defined to constrain these quantities. “Option C” could meet all the SAM-CAAM objectives with about 20 instruments, most of which have flown before, but never routinely several times per week, and never as a group. Aircraft integration and approaches to data handling, payload support, and logistical considerations for a long-term, operational mission are discussed. Finally, SAM-CAAM is feasible because, for most aerosol sources and specified seasons, particle properties tend to be repeatable, even if aerosol loading varies.« less

neoPASCAL: A Cubesat-based approach to validate Mars GCMs using a network of landed sensors

NASA Astrophysics Data System (ADS)

Moores, John; Podmore, Hugh; Lee, Regina S. K.; Haberle, Robert

2017-10-01

Beginning in the 1990s, concepts for a network of 15-20 small (12.8 kg) landers to measure surface pressure across Mars were proposed (Merrihew et al., 1996). Such distributed measurements were seen as particularly valuable as they held the promise of validating Mars Global Circulation Models (GCMs), for which the diurnal and seasonal variations in surface pressure may be diagnostically related to atmospheric parameters (Haberle et al., 1996). MicroMET, later renamed PASCAL, was a Discovery contender, however, the total mass required for the 20 landers and a support orbiter presented a challenge compared to the delivered science.In the 20 years since this concept originated, miniaturization of spacecraft systems, sensors and components has made substantial progress. Several small planetary science spacecraft based on the CubeSat design approach will launch in the next few years. Yet, only one meteorological station (REMS) currently operates on the surface of Mars. Meanwhile, the output from atmospheric models have become ever more critical for understanding key Martian geological processes including volatile transport, identifying the extent and persistence of surface brines, understanding the sources and sinks of methane and investigating the past climate of Mars, to name only a few areas.As such, it is time to reconsider the PASCAL concept. We find that modern equipment opens up payload space in the original 12.8 kg entry-vehicles from 23 g to nearly 1 kg, sufficient for adding small imagers, spectrometers and other additional or alternate payloads to examine atmosphere and surface over a wide geographic range of settings. If, instead, we seek the minimum solution for spacecraft mass, we find that a pressure-sensing vehicle would mass < 250 g at entry making these spacecraft appealing secondary payloads for future Mars missions.

Future space transportation systems analysis study. Phase 1 extension: Transportation systems reference data, volume 2

NASA Technical Reports Server (NTRS)

1975-01-01

Transportation mass requirements are developed for various mission and transportation modes based on vehicle systems sized to fit the exact needs of each mission. The parametric data used to derive the mass requirements for each mission and transportation mode are presented to enable accommodation of possible changes in mode options or payload definitions. The vehicle sizing and functional requirements used to derive the parametric data are described.

Antares Rocket Test Launch

NASA Image and Video Library

2013-04-21

The Orbital Sciences Corporation Antares rocket is seen as it launches from Pad-0A of the Mid-Atlantic Regional Spaceport (MARS) at the NASA Wallops Flight Facility in Virginia, Sunday, April 21, 2013. The test launch marked the first flight of Antares and the first rocket launch from Pad-0A. The Antares rocket delivered the equivalent mass of a spacecraft, a so-called mass simulated payload, into Earth's orbit. Photo Credit: (NASA/Bill Ingalls)

Payload and Survivability Tradeoffs in the Presence of Risk

DTIC Science & Technology

2013-06-01

9. SPONSORING /MONITORING AGENCY NAME(S) AND ADDRESS(ES) N/ A 10. SPONSORING/MONITORING AGENCY REPORT NUMBER 11. SUPPLEMENTARY NOTES The...DISTRIBUTION CODE A 13. ABSTRACT (maximum 200 words) The objective of this thesis research is to find an optimized throughput plan for ship...considering cargo ship size, quantity, speed, range, and risk when traversing through a designated travel area. Data collected from previous theses

UltraSail - Ultra-Lightweight Solar Sail Concept

NASA Technical Reports Server (NTRS)

Burton, Rodney L.; Coverstone, Victoria L.; Hargens-Rysanek, Jennifer; Ertmer, Kevin M.; Botter, Thierry; Benavides, Gabriel; Woo, Byoungsam; Carroll, David L.; Gierow, Paul A.; Farmer, Greg

2005-01-01

UltraSail is a next-generation high-risk, high-payoff sail system for the launch, deployment, stabilization and control of very large (sq km class) solar sails enabling high payload mass fractions for high (Delta)V. Ultrasail is an innovative, non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation-flying micro-satellites with an innovative solar sail architecture to achieve controllable sail areas approaching 1 sq km, sail subsystem area densities approaching 1 g/sq m, and thrust levels many times those of ion thrusters used for comparable deep space missions. Ultrasail can achieve outer planetary rendezvous, a deep space capability now reserved for high-mass nuclear and chemical systems. One of the primary innovations is the near-elimination of sail supporting structures by attaching each blade tip to a formation-flying micro-satellite which deploys the sail, and then articulates the sail to provide attitude control, including spin stabilization and precession of the spin axis. These tip micro-satellites are controlled by 3-axis micro-thruster propulsion and an on-board metrology system. It is shown that an optimum spin rate exists which maximizes payload mass.

NASA's Space Launch System: A Transformative Capability for Deep Space Missions

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2017-01-01

Already making substantial progress toward its first launches, NASA’s Space Launch System (SLS) exploration-class launch vehicle presents game-changing new opportunities in spaceflight, enabling human exploration of deep space, as well as a variety of missions and mission profiles that are currently impossible. Today, the initial configuration of SLS, able to deliver more than 70 metric tons of payload to low Earth orbit (LEO), is well into final production and testing ahead of its planned first flight, which will send NASA’s new Orion crew vehicle around the moon and will deploy 13 CubeSats, representing multiple disciplines, into deep space. At the same time, production work is already underway toward the more-capable Block 1B configuration, planned to debut on the second flight of SLS, and capable of lofting 105 tons to LEO or of co-manifesting large exploration systems with Orion on launches to the lunar vicinity. Progress being made on the vehicle for that second flight includes initial welding of its core stage and testing of one of its engines, as well as development of new elements such as the powerful Exploration Upper Stage and the Universal Stage Adapter “payload bay.” Ultimately, SLS will evolve to a configuration capable of delivering more than 130 tons to LEO to support humans missions to Mars. In order to enable human deep-space exploration, SLS provides unrivaled mass, volume, and departure energy for payloads, offering numerous benefits for a variety of other missions. For robotic science probes to the outer solar system, for example, SLS can cut transit times to less than half that of currently available vehicles or substantially increased spacecraft mass. In the field of astrophysics, SLS’ high payload volume, in the form of payload fairings with a diameter of up to 10 meters, creates the opportunity for launch of large-aperture telescopes providing an unprecedented look at our universe. This presentation will give an overview of SLS’ capabilities and its current status, and discuss the vehicle’s potential for human exploration of deep space and other game-changing utilization opportunities.

Cost-Effective NEO Characterization Using Solar Electric Propulsion (SEP)

NASA Astrophysics Data System (ADS)

Dissly, R. W.; Reinert, R.; Mitchell, S.

2003-05-01

We present a cost-effective multiple NEO rendezvous mission design optimized around the capabilities of Ball's 200-kg NEOX Solar Electric Propelled microsatellite. The NEOX spacecraft is 3-axis stabilized with better-than 1 milliradian pointing accuracy to serve as an excellent imaging platform; its DSN compatible telecommunications subsystem can support a 6.4-kbps downlink rate at 3 AU earth range. The spacecraft mass is <200kg at launch to allow launch as a cost-effective secondary payload. It uses proven SEP technology to provide 12km/s of Delta-V, which enables multiple rendezvous' in a single mission. Cost-effectiveness is optimized by launch as a secondary payload (e.g., Ariane-5 ASAP) or as a multiple manifest on a single dedicated launch vehicle (e.g., 4 on a Delta-II 2925). Following separation from the LV, we describe a candidate mission profile that minimizes cost by using the spacecraft's 12km/s of SEP Delta-V to allow orbiting up to 4 separate NEO's. Orbiting as opposed to flying by augments the mission's science return by providing the NEO mass and by allowing multiple phase angle imaging. The NEOX Spacecraft has the capability to support a 20kg payload drawing 100W average during SEP cruise, with >1kW available during the NEO orbital phase when the SEP thrusters are not powered. We will present a candidate payload suite that includes a visible/NIR imager, a laser altimeter, and a set of small, self-righting surface probes that can be used to assess the geophysical state of the object surface and near-surface environments. The surface probe payload notionally includes a set of cameras for imaging the body surface at mm-scale resolution, an accelerometer package to measure surface mechanical properties upon probe impact, a Langmuir probe to measure the electrostatic gradient immediately above the object surface, and an explosive charge that can be remotely detonated at the end of the surface mission to excavate an artificial crater that can be remotely observed from the orbiting spacecraft.

RD860 and RD860L Engines with Deep Thrust Throttling and a High Technology Readiness Level (TRL)

NASA Astrophysics Data System (ADS)

Prokopchuk, O. O.; Shul'ga, V. A.; Dibrivnyi, O. V.; Kukhta, A. S.

2018-04-01

To solve the problems of delivering payloads to Mars surface and returning them to the orbit, liquid rocket engines, operating on storable propellants with deep throttling possibility, are needed, besides having high energy-mass characteristics.

Places Only Sails Can Go

NASA Technical Reports Server (NTRS)

Montgomery, Edward E., IV; Heaton, Andrew F.; Garbe, Gregory P.

2003-01-01

Solar sails are a near term, low thrust, propellantless propulsion technology suitable for orbital maneuvering, station keeping, and attitude control applications for small payloads. Furthermore, these functions can be highly integrated, reducing mass, cost and complexity. The solar sail concept is based on momentum exchange with solar flux reflected from a large, deployed thin membrane. Thrust performance increases as the square of the distance to the sun. In comparison to conventional chemical systems, there are missions where solar sails are vastly more and less economical. The less attractive applications involve large payloads, outer solar system transfers, and short trip times. However, for inclination changes and station keeping at locations requiring constant thrust, the solar sail is the only economical option for missions of more than a few weeks duration. We compare the location and energies required for these applications between solar sails, advanced electric propulsion, and conventional rockets. We address the effect on mass fraction to understand solar sail mission cost and capability. Finally, the benefit of potential applications to near term science missions is reported.

Aerocapture Inflatable Decelerator (AID)

NASA Technical Reports Server (NTRS)

Reza, Sajjad

2007-01-01

Forward Attached Inflatable Decelerators, more commonly known as inflatable aeroshells, provide an effective, cost efficient means of decelerating spacecrafts by using atmospheric drag for aerocapture or planetary entry instead of conventional liquid propulsion deceleration systems. Entry into planetary atmospheres results in significant heating and aerodynamic pressures which stress aeroshell systems to their useful limits. Incorporation of lightweight inflatable decelerator surfaces with increased surface-area footprints provides the opportunity to reduce heat flux and induced temperatures, while increasing the payload mass fraction. Furthermore, inflatable aeroshell decelerators provide the needed deceleration at considerably higher altitudes and Mach numbers when compared with conventional rigid aeroshell entry systems. Inflatable aeroshells also provide for stowage in a compact space, with subsequent deployment of a large-area, lightweight heatshield to survive entry heating. Use of a deployable heatshield decelerator not only enables an increase in the spacecraft payload mass fraction and but may also eliminate the need for a spacecraft backshell and cruise stage. This document is the viewgraph slides for the paper's presentation.

CNES-NASA Studies of the Mars Sample Return Orbiter Aerocapture Phase

NASA Technical Reports Server (NTRS)

Fraysse, H.; Powell, R.; Rousseau, S.; Striepe, S.

2000-01-01

A Mars Sample Return (MSR) mission has been proposed as a joint CNES (Centre National d'Etudes Spatiales) and NASA effort in the ongoing Mars Exploration Program. The MSR mission is designed to return the first samples of Martian soil to Earth. The primary elements of the mission are a lander, rover, ascent vehicle, orbiter, and an Earth entry vehicle. The Orbiter has been allocated only 2700 kg on the launch phase to perform its part of the mission. This mass restriction has led to the decision to use an aerocapture maneuver at Mars for the orbiter. Aerocapture replaces the initial propulsive capture maneuver with a single atmospheric pass. This atmospheric pass will result in the proper apoapsis, but a periapsis raise maneuver is required at the first apoapsis. The use of aerocapture reduces the total mass requirement by approx. 45% for the same payload. This mission will be the first to use the aerocapture technique. Because the spacecraft is flying through the atmosphere, guidance algorithms must be developed that will autonomously provide the proper commands to reach the desired orbit while not violating any of the design parameters (e.g. maximum deceleration, maximum heating rate, etc.). The guidance algorithm must be robust enough to account for uncertainties in delivery states, atmospheric conditions, mass properties, control system performance, and aerodynamics. To study this very critical phase of the mission, a joint CNES-NASA technical working group has been formed. This group is composed of atmospheric trajectory specialists from CNES, NASA Langley Research Center and NASA Johnson Space Center. This working group is tasked with developing and testing guidance algorithms, as well as cross-validating CNES and NASA flight simulators for the Mars atmospheric entry phase of this mission. The final result will be a recommendation to CNES on the algorithm to use, and an evaluation of the flight risks associated with the algorithm. This paper will describe the aerocapture phase of the MSR mission, the main principles of the guidance algorithms that are under development, the atmospheric entry simulators developed for the evaluations, the process for the evaluations, and preliminary results from the evaluations.

The Drop Tower Bremen -Experiment Operation

NASA Astrophysics Data System (ADS)

Könemann, Thorben; von Kampen, Peter; Rath, Hans J.

The idea behind the drop tower facility of the Center of Applied Space Technology and Micro-gravity (ZARM) in Bremen is to provide an inimitable technical opportunity of a daily access to short-term weightlessness on earth. In this way ZARM`s european unique ground-based microgravity laboratory displays an excellent economic alternative for research in space-related conditions at low costs comparable to orbital platforms. Many national and international ex-perimentalists motivated by these prospects decide to benefit from the high-quality and easy accessible microgravity environment only provided by the Drop Tower Bremen. Corresponding experiments in reduced gravity could open new perspectives of investigation methods and give scientists an impressive potential for a future technology and multidisciplinary applications on different research fields like Fundamental Physics, Astrophysics, Fluid Dynamics, Combus-tion, Material Science, Chemistry and Biology. Generally, realizing microgravity experiments at ZARM`s drop tower facility meet new requirements of the experimental hardware and may lead to some technical constraints in the setups. In any case the ZARM Drop Tower Operation and Service Company (ZARM FAB mbH) maintaining the drop tower facility is prepared to as-sist experimentalists by offering own air-conditioned laboratories, clean rooms, workshops and consulting engineers, as well as scientific personal. Furthermore, ZARM`s on-site apartment can be used for accommodations during the experiment campaigns. In terms of approaching drop tower experimenting, consulting of experimentalists is mandatory to successfully accomplish the pursued drop or catapult capsule experiment. For this purpose there will be a lot of expertise and help given by ZARM FAB mbH in strong cooperation to-gether with the experimentalists. However, in comparison to standard laboratory setups the drop or catapult capsule setup seems to be completely different at first view. While defining a microgravity project at the Drop Tower Bremen, interesting experimentalists should keep in mind generally reducing dimensions and masses of their common laboratory setups to meet the capsule constraints: overall payload height 980mm/1730mm (short/long drop capsule) and 950mm (catapult capsule); area of each capsule platform 0,359sqm; maximum payload mass 274kg/234kg (short/long drop capsule) and 163,8kg (catapult capsule). The base equipments of each capsule are the Capsule Control System (CCS) to remote control the experiment and the rechargeable battery pack (24V/40A) for the experiment operation. Moreover, the exper-iment components must be able to withstand maximum decelerations of 50g while the short capsule impact of about 200ms, and maximum accelerations of 30g while catapult launch with a duration of about 300ms. In our second talk concerning ZARM`s drop tower facility we will go on with detailed infor-mations about the technical base setups of the drop and the catapult capsule structure to completely handle a freely falling experiment. Furthermore, we will summarize interesting current drop tower projects as an outlook to present you the range of opportunities at the ground-based short-term microgravity laboratory of ZARM.

SMART-1/CLEMENTINE Study of Humorum and Procellarum Basins

NASA Astrophysics Data System (ADS)

Carey, William; Foing, Bernard H.; Koschny, Detlef; Pio Rossi, Angelo; Josset, Jean-Luc

A study undertaken by ESA to define a European Reference Architecture for Space Exploration is due to be completed in September 2008. The development of this architecture over the past twelve months has identified a number of key capabilities, among them a lunar lander system, which could form the basis for Europe's contribution to the future exploration of space in collaboration with International Partners. The focus of this paper will be on the lunar lander system, and will present the results of an analysis of possible payloads that could be accommodated by the lander. As the industrial study is at the Phase 0 or Pre-Phase A level, the design of such a lander system is at a very early stage in its development, but an estimation of the payload capacity allows a general assessment of the types of possible payloads that could be carried, currently this capacity is estimated at 1.1 tonnes of gross payload mass to the lunar surface (assuming an Ariane 5 ECA launch). An important characteristic of the lunar lander is that it provides a versatile and flexible system for utilisation in a broad range of lunar missions which include: - Independent lunar exploration missions for science, technology demonstration and research. - Delivery of logistics and cargo to support human surface sortie missions. - Delivery of logistics to a lunar base/outpost. - Deployment of individual infrastructure elements in support of a lunar base/outpost. Based on the above different types of missions, a number of configurations of "reference payload" sets are in the process of being defined that cover specific exploration objectives related primarily to capability demonstration, exploration enabling research and enabled science. Aspects covered include: ISRU, robotics, mobility, human preparation, life science and geology. This paper will present the current status of definition of the Reference Payload sets.

Human Mars Mission: SEP Architecture Crew Taxi Propulsion Stage Study and Design and Technology for Reaction and Control System. Part 1

NASA Technical Reports Server (NTRS)

Young, Archie

1999-01-01

The Mars exploration is a candidate pathway to expand human presence and useful activities in the solar system. There are several propulsion system options being considered to place the Mars payload on its inter-planetary transfer trajectory. One propulsion option is the use of Solar Electric Propulsion (SEP) to spiral out with the Mars payload from an initial Low Earth Orbit (LEO) to an elliptical High Earth Orbit (HEO). This report, presented in annotated facing page format, describes the work completed on the design of a crew taxi propulsion stage used in conjunction with the SEP. Transportation system/mission analysis topics covered in this report include sub-system analysis, trajectory profile description, mass performance and crew taxi stage sizing, stage configuration, stage cost, and Trans-Mars Injection (TMI) launch window. The high efficiency of SEP is used to provide the major part of the TMI propulsion maneuver. Orbital energy is continuously added over a period of approximately twelve months. The SEP and Mars payload follow a spiral trajectory from an initial LEO to a final elliptical HEO. A small chemical stage is then used to provide the final part of the TMI. The now unloaded SEP returns to LEO to repeat another spiral trajectory with payload to HEO. The spiral phase of the SEP's trajectory takes several months to reach HEO, thus significantly increasing the exposure time of the crew to zero-gravity. In order to minimize the long zero-gravity effects, a high thrust chemical stage delivers the crew to the SEP's HEO. The crew rendezvous with the Mars payload in HEO. After a checkout period the Mars payload with the crew is injected onto a Trans-Mars Trajectory by a small chemical stage.

Human Mars Mission: SEP Architecture, Crew Taxi Propulsion Stage Study and Design and Technology for Reaction and Control System. Pt. 1

NASA Technical Reports Server (NTRS)

Young, Archie

1999-01-01

The Mars exploration is a candidate pathway to expand human presence and useful activities in the solar system. There are several propulsion system options being considered to place the Mars payload on its interplanetary transfer trajectory. One propulsion option is the use of Solar Electric Propulsion (SEP) to spiral out with the Mars payload from an initial Low Earth Orbit (LEO) to an elliptical High Earth Orbit (HEO). This report, presented in annotated facing page format, describes the work completed on the design of a crew taxi propulsion stage used in conjunction with the SEP. Transportation system/mission analysis topics covered in this report include sub-system analysis, trajectory profile description, mass performance and crew taxi stage sizing, stage configuration, stage cost, and Trans-Mars Injection (TMI) launch window. The high efficiency of SEP is used to provide the major part of the TMI propulsion maneuver. Orbital energy is continuously added over a period of approximately twelve months. The SEP and Mars payload follow a spiral trajectory from an initial LEO to a final elliptical HEO. A small chemical stage is then used to provide the final part of the TMI. The now unloaded SEP returns to LEO to repeat another spiral trajectory with payload to HEO. The spiral phase of the SEP's trajectory takes several months to reach HEO, thus significantly increasing the exposure time of the crew to zero-gravity. In order to minimize the long zero-gravity effects, a high thrust chemical stage delivers the crew to the SEP's HEO. The crew rendezvous with the Mars payload in HEO. After a checkout period the Mars payload with the crew is injected onto a Trans-Mars Trajectory by a small chemical stage.

KSC-00pp1737

NASA Image and Video Library

2000-11-14

Workers in the payload changeout room stand by as the doors open on the payload transport canister. Inside is the P6 integrated truss segment, which will fly on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

Ares V: New Opportunities for Scientific Payloads

NASA Technical Reports Server (NTRS)

Creech, Steve

2009-01-01

What if scientists and payload planners had access to three to five times the volume and five to nine times the mass provided by today's launch vehicles? This simple question can lead to numerous exciting possibilities, all involving NASA's new Ares V cargo launch vehicle now on the drawing board. Multiple scientific fields and payload designers have that opportunity with the Ares V cargo launch vehicle, being developed at NASA as the heavy-lift component of the U.S. Space Exploration Policy. When the Ares V begins flying late next decade, its capabilities will significantly exceed the 1960s-era Saturn V or the current Space Shuttle, while it benefits from their engineering, manufacturing, and infrastructure heritage. It will send more crew and cargo to more places on the lunar surface than Apollo and provide ongoing support to a permanent lunar outpost. Moreover, it will restore a strategic heavy-lift U.S. asset, which can support human and robotic exploration and scientific ventures for decades to come. Assessment of astronomy payload requirements since Spring 2008 has indicated that Ares V has the potential to support a range of payloads and missions. Some of these missions were impossible in the absence of Ares V's capabilities. Collaborative design/architecture inputs, exchanges, and analyses have already begun between scientists and payload developers. A 2008 study by a National Research Council (NRC) panel, as well as analyses presented by astronomers and planetary scientists at two weekend conferences in 2008, support the position that Ares V has benefit to a broad range of planetary and astronomy missions. This early dialogue with Ares V engineers is permitting the greatest opportunity for payload/transportation/mission synergy and the least financial impact to Ares V development. In addition, independent analyses suggest that Ares V has the opportunity to enable more cost-effective mission design.

Sparsity prediction and application to a new steganographic technique

NASA Astrophysics Data System (ADS)

Phillips, David; Noonan, Joseph

2004-10-01

Steganography is a technique of embedding information in innocuous data such that only the innocent data is visible. The wavelet transform lends itself to image steganography because it generates a large number of coefficients representing the information in the image. Altering a small set of these coefficients allows embedding of information (payload) into an image (cover) without noticeably altering the original image. We propose a novel, dual-wavelet steganographic technique, using transforms selected such that the transform of the cover image has low sparsity, while the payload transform has high sparsity. Maximizing the sparsity of the payload transform reduces the amount of information embedded in the cover, and minimizing the sparsity of the cover increases the locations that can be altered without significantly altering the image. Making this system effective on any given image pair requires a metric to indicate the best (maximum sparsity) and worst (minimum sparsity) wavelet transforms to use. This paper develops the first stage of this metric, which can predict, averaged across many wavelet families, which of two images will have a higher sparsity. A prototype implementation of the dual-wavelet system as a proof of concept is also developed.

A Unique Power System For The ISS Fluids And Combustion Facility

NASA Technical Reports Server (NTRS)

Fox, David A.; Poljak, Mark D.

2001-01-01

Unique power control technology has been incorporated into an electrical power control unit (EPCU) for the Fluids and Combustion Facility (FCF). The objective is to maximize science throughput by providing a flexible power system that is easily reconfigured by the science payload. Electrical power is at a premium on the International Space Station (ISS). The EPCU utilizes advanced power management techniques to maximize the power available to the FCF experiments. The EPCU architecture enables dynamic allocation of power from two ISS power channels for experiments. Because of the unique flexible remote power controller (FRPC) design, power channels can be paralleled while maintaining balanced load sharing between the channels. With an integrated and redundant architecture, the EPCU can tolerate multiple faults and still maintain FCF operation. It is important to take full advantage of the EPCU functionality. The EPCU acts as a buffer between the experimenter and the ISS power system with all its complex requirements. However, FCF science payload developers will still need to follow guidelines when designing the FCF payload power system. This is necessary to ensure power system stability, fault coordination, electromagnetic compatibility, and maximum use of available power for gathering scientific data.

Predicted Performance of a Thrust-Enhanced SR-71 Aircraft with an External Payload

NASA Technical Reports Server (NTRS)

Conners, Timothy R.

1997-01-01

NASA Dryden Flight Research Center has completed a preliminary performance analysis of the SR-71 aircraft for use as a launch platform for high-speed research vehicles and for carrying captive experimental packages to high altitude and Mach number conditions. Externally mounted research platforms can significantly increase drag, limiting test time and, in extreme cases, prohibiting penetration through the high-drag, transonic flight regime. To provide supplemental SR-71 acceleration, methods have been developed that could increase the thrust of the J58 turbojet engines. These methods include temperature and speed increases and augmentor nitrous oxide injection. The thrust-enhanced engines would allow the SR-71 aircraft to carry higher drag research platforms than it could without enhancement. This paper presents predicted SR-71 performance with and without enhanced engines. A modified climb-dive technique is shown to reduce fuel consumption when flying through the transonic flight regime with a large external payload. Estimates are included of the maximum platform drag profiles with which the aircraft could still complete a high-speed research mission. In this case, enhancement was found to increase the SR-71 payload drag capability by 25 percent. The thrust enhancement techniques and performance prediction methodology are described.

KSC-08pd2656

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, technicians finish replacing the protective cover over the Super Lightweight Interchangeable Carrier, or SLIC. The cover was removed to clean the carrier of contaminants found Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2643

NASA Image and Video Library

2008-09-17

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, a technician uncovers the Super Lightweight Interchangeable Carrier, or SLIC. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A will be removed. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies for servicing of the Hubble Space Telescope on the STS-125 mission, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2646

NASA Image and Video Library

2008-09-17

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, technicians inspect areas of the Super Lightweight Interchangeable Carrier, or SLIC, for contamination. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A will be removed. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies for servicing of the Hubble Space Telescope on the STS-125 mission, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2653

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, the protective cover is being replaced on the Super Lightweight Interchangeable Carrier, or SLIC. The cover was removed to clean the carrier of contaminants found Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2645

NASA Image and Video Library

2008-09-17

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, technicians inspect areas of the Super Lightweight Interchangeable Carrier, or SLIC, for contamination. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A will be removed. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies for servicing of the Hubble Space Telescope on the STS-125 mission, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2655

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, a technician works to replace the protective cover on the Super Lightweight Interchangeable Carrier, or SLIC. The cover was removed to clean the carrier of contaminants found Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2654

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, a technician works to replace the protective cover on the Super Lightweight Interchangeable Carrier, or SLIC. The cover was removed to clean the carrier of contaminants found Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2657

NASA Image and Video Library

2008-09-18

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, technicians finish replacing the protective cover over the Super Lightweight Interchangeable Carrier, or SLIC. The cover was removed to clean the carrier of contaminants found Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2642

NASA Image and Video Library

2008-09-17

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, a technician begins uncovering the Super Lightweight Interchangeable Carrier, or SLIC. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A will be removed. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies for servicing of the Hubble Space Telescope on the STS-125 mission, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

KSC-08pd2644

NASA Image and Video Library

2008-09-17

CAPE CANAVERAL, Fla. - In the Payload Hazardous Servicing Facility at NASA's Kennedy Space Center, technicians inspect areas of the Super Lightweight Interchangeable Carrier, or SLIC, for contamination. Contamination discovered Sept. 17 during preparations to deliver NASA's Hubble Space Telescope servicing payload to Launch Pad 39A will be removed. Cleanliness is extremely important for space shuttle Atlantis’ STS-125 mission to Hubble, and the teams have insured that the SLIC is ready to fly. The SLIC, which holds battery module assemblies for servicing of the Hubble Space Telescope on the STS-125 mission, is built with state-of-the-art, lightweight, composite materials - carbon fiber with a cyanate ester resin and a titanium metal matrix composite. These composites have greater strength-to-mass ratios than the metals typically used in spacecraft design. The carrier is one of four being transferred to Launch Pad 39A. At the pad, the carriers will be loaded into Atlantis’ payload bay. Launch of Atlantis is targeted for Oct. 10. Photo credit: NASA/Jack Pfaller

The first dedicated life sciences mission - Spacelab 4

NASA Technical Reports Server (NTRS)

Cramer, D. R.; Reid, D. H.; Klein, H. P.

1983-01-01

The details of the payload and the experiments in Spacelab 4, the first Spacelab mission dedicated entirely to the life sciences, are discussed. The payload of Spacelab 4, carried in the bay of the Shuttle Orbiter, consists of 25 tentatively selected investigations combined into a comprehensive integrated exploration of the effects of acute weightlessness on living systems. The payload contains complementary designs in the human and animal investigations in order to validate animal models of human physiology in weightlessness. Animals used as experimental subjects will include squirrel monkeys, laboratory rats, several species of plants, and frog eggs. The main scientific objectives of the investigations include the study of the acute cephalic fluid shift, cardiovascular adaptation to weightlessness, including postflight reductions in orthostatic tolerance and exercise capacity, and changes in vestibular function, including space motion sickness, associated with weightlessness. Other scientific objective include the study of red cell mass reduction, negative nitrogen balance, altered calcium metabolism, suppressed in vitro lymphocyte reactivity, gravitropism and photropism in plants, and fertilization and early development in frog eggs.

NEXT-Lunar Lander -an Opportunity for a Close Look at the Lunar South Pole

NASA Astrophysics Data System (ADS)

Homeister, Maren; Thaeter, Joachim; Scheper, Marc; Apeldoorn, Jeffrey; Koebel, David

The NEXT-Lunar Lander mission, as contracted by ESA and investigated by OHB-System and its industrial study team, has two main purposes. The first is technology demonstration for enabling technologies like propulsion-based soft precision landing for future planetary landing missions. This involves also enabling technology experiments, like fuel cell, life science and life support, which are embedded in the stationary payload of the lander. The second main and equally important aspect is the in-situ investigation of the surface of the Moon at the lunar South Pole by stationary payload inside the Lander, deployable payload to be placed in the vicinity of the lander and mobile payload carried by a rover. The currently assessed model payload includes 15 instruments on the lander and additional five on the rover. They are addressing the fields geophysics, geochemistry, geology and radio astronomy preparation. The mission is currently under investigation in frame of a phase A mission study contract awarded by ESA to two independent industrial teams, of which one is led by OHB-System. The phase A activities started in spring 2008 and were conducted until spring 2010. A phase B is expected shortly afterwards. The analysed mission architectures range from a Soyuz-based mission to a Shared-Ariane V class mission via different transfer trajectories. Depending on the scenario payload masses including servicing of 70 to 150 kg can be delivered to the lunar surface. The lander can offer different services to the payload. The stationary payload is powered and conditioned by the lander. Examples for embarked payloads are an optical camera system, a Radio Science Experiment and a radiation monitor. The lander surface payload is deployed to the lunar surface by a 5 DoF robotic arm and will be powered by the Lander. To this group of payloads belong seismometers, a magnetometer and an instrumented Mole. The mobile payload will be carried by a rover. The rover is equipped with its own 5 DoF robotic arm and can travel with an average speed of about 1 cm/s. The Rover is generally tele-operated but has the capability to execute autonomously pre-selected operation tasks, is aware of its current status and analyses potential hazards to avoid loss of its mission by operator failure. It is equipped with a model payload consisting of a camera system for multi-spectra including infra-red, a Raman-LIBS and a CLUPI. In addition its task is to position seismometers at a distance of about 1 km away from the lander. The baseline scenario includes a launch in the 2018 timeframe and one year of surface operations at the Shakleton crater rim. This presentation will focus on the following points: • Mission architecture and spacecraft layout as elaborated during the past study activities • Surface operations of lander and rover • Current mission capability to support scientific investigations at the lunar South Pole

Mars MetNet Mission - Martian Atmospheric Observational Post Network

NASA Astrophysics Data System (ADS)

Harri, Ari-Matti; Aleksashkin, Sergey; Arruego, Ignacio; Schmidt, Walter; Ponomarenko, Andrey; Apestigue, Victor; Genzer, Maria; Vazquez, Luis; Uspensky, Mikhail; Haukka, Harri

2016-04-01

A new kind of planetary exploration mission for Mars is under development in collaboration between the Finnish Meteorological Institute (FMI), Lavochkin Association (LA), Space Research Institute (IKI) and Institutio Nacional de Tecnica Aerospacial (INTA). The Mars MetNet mission is based on a new semi-hard landing vehicle called MetNet Lander (MNL). The scientific payload of the Mars MetNet Precursor [1] mission is divided into three categories: Atmospheric instruments, Optical devices and Composition and structure devices. Each of the payload instruments will provide significant insights in to the Martian atmospheric behavior. The key technologies of the MetNet Lander have been qualified and the electrical qualification model (EQM) of the payload bay has been built and successfully tested. MetNet Lander The MetNet landing vehicles are using an inflatable entry and descent system instead of rigid heat shields and parachutes as earlier semi-hard landing devices have used. This way the ratio of the payload mass to the overall mass is optimized. The landing impact will burrow the payload container into the Martian soil providing a more favorable thermal environment for the electronics and a suitable orientation of the telescopic boom with external sensors and the radio link antenna. It is planned to deploy several tens of MNLs on the Martian surface operating at least partly at the same time to allow meteorological network science. Strawman Scientific Payload The strawman payload of the two MNL precursor models includes the following instruments: Atmospheric instruments: • MetBaro Pressure device • MetHumi Humidity device • MetTemp Temperature sensors Optical devices: • PanCam Panoramic • MetSIS Solar irradiance sensor with OWLS optical wireless system for data transfer • DS Dust sensor Composition and Structure Devices: • Tri-axial magnetometer MOURA • Tri-axial System Accelerometer The descent processes dynamic properties are monitored by a special 3-axis accelerometer combined with a 3-axis gyrometer. The data will be sent via auxiliary beacon antenna throughout the descent phase starting shortly after separation from the spacecraft. MetNet Mission payload instruments are specially designed to operate under very low power conditions. MNL flexible solar panels provides a total of approximately 0.7-0.8 W of electric power during the daylight time. As the provided power output is insufficient to operate all instruments simultaneously they are activated sequentially according to a specially designed cyclogram table which adapts itself to the different environmental constraints. Mission Status Full Qualification Model (QM) of the MetNet landing unit with the Precursor Mission payload is currently under functional tests. In the near future the QM unit will be exposed to environmental tests with qualification levels including vibrations, thermal balance, thermal cycling and mechanical impact shock. One complete flight unit of the entry, descent and landing systems (EDLS) has been manufactured and tested with acceptance levels. Another flight-like EDLS has been exposed to most of the qualification tests, and hence it may be used for flight after refurbishments. Accordingly two flight-capable EDLS systems exist. The eventual goal is to create a network of atmospheric observational posts around the Martian surface. Even if the MetNet mission is focused on the atmospheric science, the mission payload will also include additional kinds of geophysical instrumentation. The next step in the MetNet Precursor Mission is the demonstration of the technical robustness and scientific capabilities of the MetNet type of landing vehicle. Definition of the Precursor Mission and discussions on launch opportunities are currently under way. The baseline program development funding exists for the next five years. Flight unit manufacture of the payload bay takes about 18 months, and it will be commenced after the Precursor Mission has been defined. References [1] http://metnet.fmi.fi

Mars MetNet Mission Status

NASA Astrophysics Data System (ADS)

Harri, Ari-Matti; Aleksashkin, Sergei; Arruego, Ignacio; Schmidt, Walter; Genzer, Maria; Vazquez, Luis; Haukka, Harri

2015-04-01

New kind of planetary exploration mission for Mars is under development in collaboration between the Finnish Meteorological Institute (FMI), Lavochkin Association (LA), Space Research Institute (IKI) and Institutio Nacional de Tecnica Aerospacial (INTA). The Mars MetNet mission is based on a new semi-hard landing vehicle called MetNet Lander (MNL). The scientific payload of the Mars MetNet Precursor [1] mission is divided into three categories: Atmospheric instruments, Optical devices and Composition and structure devices. Each of the payload instruments will provide significant insights in to the Martian atmospheric behavior. The key technologies of the MetNet Lander have been qualified and the electrical qualification model (EQM) of the payload bay has been built and successfully tested. 1. MetNet Lander The MetNet landing vehicles are using an inflatable entry and descent system instead of rigid heat shields and parachutes as earlier semi-hard landing devices have used. This way the ratio of the payload mass to the overall mass is optimized. The landing impact will burrow the payload container into the Martian soil providing a more favorable thermal environment for the electronics and a suitable orientation of the telescopic boom with external sensors and the radio link antenna. It is planned to deploy several tens of MNLs on the Martian surface operating at least partly at the same time to allow meteorological network science. 2. Scientific Payload The payload of the two MNL precursor models includes the following instruments: Atmospheric instruments: 1. MetBaro Pressure device 2. MetHumi Humidity device 3. MetTemp Temperature sensors Optical devices: 1. PanCam Panoramic 2. MetSIS Solar irradiance sensor with OWLS optical wireless system for data transfer 3. DS Dust sensor The descent processes dynamic properties are monitored by a special 3-axis accelerometer combined with a 3-axis gyrometer. The data will be sent via auxiliary beacon antenna throughout the descent phase starting shortly after separation from the spacecraft. MetNet Mission payload instruments are specially designed to operate in very low power conditions. MNL flexible solar panels provides a total of approximately 0.7-0.8 W of electric power during the daylight time. As the provided power output is insufficient to operate all instruments simultaneously they are activated sequentially according to a specially designed cyclogram table which adapts itself to the different environmental constraints. 3. Mission Status Full Qualification Model (QM) of the MetNet landing unit with the Precursor Mission payload is currently under functional tests. In near future the QM unit will be exposed to environmental tests with qualification levels including vibrations, thermal balance, thermal cycling and mechanical impact shock. One complete flight unit of the entry, descent and landing systems (EDLS) has been manufactured and tested with acceptance levels. Another flight-like EDLS has been exposed to most of the qualification tests, and hence it may be used for flight after refurbishments. Accordingly two flight-capable EDLS systems exist. The eventual goal is to create a network of atmospheric observational posts around the Martian surface. Even if the MetNet mission is focused on the atmospheric science, the mission payload will also include additional kinds of geophysical instrumentation. The next step in the MetNet Precursor Mission to demonstrate the technical robustness and scientific capabilities of the MetNet type of landing vehicle. Definition of the Precursor Mission and discussions on launch opportunities are currently under way. The baseline program development funding exists for the next five years. Flight unit manufacture of the payload bay takes about 18 months, and it will be commenced after the Precursor Mission has been defined. References [1] http://metnet.fmi.fi

Antares Rocket Test Launch

NASA Image and Video Library

2013-04-21

NASA Deputy Administrator Lori Garver and other guests react after having watched the successful launch of the Orbital Sciences Corporation Antares rocket from the Mid-Atlantic Regional Spaceport (MARS) at the NASA Wallops Flight Facility in Virginia, Sunday, April 21, 2013. The test launch marked the first flight of Antares and the first rocket launch from Pad-0A. The Antares rocket delivered the equivalent mass of a spacecraft, a so-called mass simulated payload, into Earth's orbit. Photo Credit: (NASA/Bill Ingalls)

LOx/LCH4: A Unifying Technology for Future Exploration

NASA Technical Reports Server (NTRS)

Banker, Brian; Ryan, Abigail

2014-01-01

OVERVIEW For every pound of payload landed on Mars, 226 pounds are required on Earth to get it there. Due to this enormous mass gear-ratio, increasing commonality between lander subsystems, such as power, propulsion, and life support, results in tremendous launch mass and cost savings. Human-Mars architectures point to an oxygen-methane economy, utilizing common commodities scavenged from the planetary atmosphere and soil via In-Situ Resource Utilization (ISRU) and common commodity tankage across sub-systems.

Volume and Mass Estimation of Three-Phase High Power Transformers for Space Applications

NASA Technical Reports Server (NTRS)

Kimnach, Greg L.

2004-01-01

Spacecraft historically have had sub-1kW(sub e), electrical requirements for GN&C, science, and communications: Galileo at 600W(sub e), and Cassini at 900W(sub e), for example. Because most missions have had the same order of magnitude power requirements, the Power Distribution Systems (PDS) use existing, space-qualified technology and are DC. As science payload and mission duration requirements increase, however, the required electrical power increases. Subsequently, this requires a change from a passive energy conversion (solar arrays and batteries) to dynamic (alternator, solar dynamic, etc.), because dynamic conversion has higher thermal and conversion efficiencies, has higher power densities, and scales more readily to higher power levels. Furthermore, increased power requirements and physical distribution lengths are best served with high-voltage, multi-phase AC to maintain distribution efficiency and minimize voltage drops. The generated AC-voltage must be stepped-up (or down) to interface with various subsystems or electrical hardware. Part of the trade-space design for AC distribution systems is volume and mass estimation of high-power transformers. The volume and mass are functions of the power rating, operating frequency, the ambient and allowable temperature rise, the types and amount of heat transfer available, the core material and shape, the required flux density in a core, the maximum current density, etc. McLyman has tabulated the performance of a number of transformers cores and derived a "cookbook" methodology to determine the volume of transformers, whereas Schawrze had derived an empirical method to estimate the mass of single-phase transformers. Based on the work of McLyman and Schwarze, it is the intent herein to derive an empirical solution to the volume and mass estimation of three-phase, laminated EI-core power transformers, having radiated and conducted heat transfer mechanisms available. Estimation of the mounting hardware, connectors, etc. is not included.

Commercial Research and Development: Power to Explore, Opportunities from Discovery

NASA Technical Reports Server (NTRS)

Casas, Joseph C.; Nall, Mark; Powers, C. Blake; Henderson, Robin N. (Technical Monitor)

2002-01-01

The technical and economic goals of commercial use of space are laudable, and are addressed as a high priority by almost every national space program and most major aerospace companies the world over. Yet, the focus of most organizational agendas and discussions tends to focus on one or two very narrow enabling aspects of this potentially large technological and economic opportunity. While government sponsored commercial launch activities and private space platforms are an integral part of efforts to leverage the commercial use of space, these activities are possibly one of the smallest parts of creating, a viable and sustainable market for the commercial use of space. Most of the current programs usually do not appropriately address some of the critical issues of the current, already interested, potential space user communities. Current programs place the focus of the majority of the user requirements on the vehicle payload weight and mass performance considerations as the primary payload economical factor in providing a commercial market with a stimulating price for gaining access to the space environment. The larger user challenges of transformation from Earth-based research and development approaches to space environment approaches are not addressed early enough in programs to impact the new business considerations of potential users. Currently, space-based research and development user activities require a large user investment in time, in development of new areas of support expertise, in development of new systems, in risk of schedule to completion, and in long term capital positioning. The larger opportunities for stimulating a strong market driven interest in commercial use of space that could result from the development of vehicle payload "leap ahead technologies" for users are being missed, and there is a real risk of limiting the potentially broader market base to support a more technologically advanced and economically lucrative outcome. A major driving force for strengthening the commercial space activities is not only the technological advances in launch vehicle, or newer satellites, but the myriad of enabling payloads technologies that could, as a goal, result in an almost transparent facilitation to regular CD a, -n access to space and microgravity environments by the future users from the existing Earth-based research and development organizations market segments. Rather than focusing only on developing high lift performance launch vehicles and then developing payloads to fit them, the real focus from a business model perspective should to be on the customer payloads requirements, and on designing launch vehicles and platforms systems for a space transportation and facility infrastructure to support all aspects of the business model for the user market. To harness the full potential of space commercialization, new efforts need to be made to comprehensively examine all the critical business model areas for commercial research, development, and manufacturing in space so as to identify specific products and efforts; to determine how such operations must be both similar to and different from current Earth-based activities; to evaluate the enabling technological devices, processes and efforts so that like efforts can be addressed in a synergistic fashion for maximum user cost effectiveness; to delineate the services that are both needed and can be provided by such activities; and to use this information to drive design and development of space commercialization efforts and policy.

STS-40 crew trains in JSC's SLS mockup located in Bldg 36

NASA Image and Video Library

1987-03-10

STS-40 Payload Specialist Millie Hughes-Fulford along with backup payload specialist Robert Ward Phillips familiarize themselves with Spacelab Life Sciences 1 (SLS-1) equipment. The two scientists are in JSC's Life Sciences Project Division (LSPD) SLS mockup located in the Bioengineering and Test Support Facility Bldg 36. Hughes-Fulford, in the center aisle, pulls equipment from an overhead stowage locker while Phillips, in the foreground, experiments with the baroreflex neck pressure chamber at Rack 11. The baroreflex collar will be used in conjuction with Experiment No. 022, Influence of Weightlessness Upon Human Autonomic Cardiovascular Control. Behind Phillips in the center aisle are body mass measurement device (BMMD) (foreground) and the stowed bicycle ergometer.

Reflective baffle for BepiColombo mission

NASA Astrophysics Data System (ADS)

Rugi-Grond, E.; Weigel, T.; Herren, A.; Dominguez Calvo, M.; Krähenbühl, U.; Mouricaud, D.; Vayssade, H.

2017-11-01

The BepiColombo Spacecraft can't tolerate to absorb a major fraction of the off-axis sunlight through larger payload apertures. Fortunately, there are solutions to design baffles such that they reflect the incoming radiation back through the front aperture rather than absorbing it. A Design Study, sponsored by ESA and performed by Contraves Space together with SAGEM Défense Securité, has analysed the potential of various solutions and assessed the options to manufacture them. The selected configuration has been analysed in detail for the optical, mechanical and thermal performance as well as the impact on mass and power dissipation. The size of the baffle was adapted to the needs of the BepiColombo Laser Altimeter (BELA) payload.

JEFF: Air transport system design simulation

NASA Technical Reports Server (NTRS)

1992-01-01

Jeff is a remotely piloted vehicle designed by the Blue Team, a division of AE441, Inc., to fulfill the mission proposed by G-Dome Enterprises: to build a cost efficient aircraft to service Aeroworld with overnight cargo delivery. The design of Jeff was most significantly influenced by the need to minimize costs. This objective was pursued by building fewer large planes as opposed to many small planes. Thus, by building an aircraft with a large payload capacity, G-Dome Enterprises will be able to minimize the large costs and the large number of cycles that are associated with a large fleet. Another factor which had a significant influence on our design was the constraint that our design had to fit into a 2'x2'x5' storage container. This constraint meant that unless we wanted to build foldable wings that Jeff's span would be limited to 10 feet. Since this was not enough lifting surface to suit our needs a canard configuration was chosen to get the needed lifting surface and avoid the structural dilemma of foldable wings. The aircraft is constructed mainly of balsa, with spruce wing and canard spars and a monokote covering. It was designed to support a maximum payload weight of 35 oz. (total aircraft weight of 108 oz.) and withstand a maximum load factor of 2.5.

The inducement of planetary boundary layer mass convergence associated with varying vorticity beneath tropospheric wind maximum

NASA Technical Reports Server (NTRS)

Johnson, D. R.

1984-01-01

The effects of the vorticity distribution are applied to study planetary boundary layer mass convergence beneath free tropospheric wind maximum. For given forcing by viscous and pressure gradient forces beneath a wind maximum, boundary layer cross stream mass transport is increased by anticyclonic vorticity on the right flank and decreased by cyclonic vorticity on the left flank. Such frictionally forced mass transport induces boundary layer mass convergence beneath the relative wind maximum. This result is related to the empirical rule that the most intense convection and severe weather frequently develop beneath the 500 mb zero relative vorticity isopleth.

Nuclear Thermal Rocket/Vehicle Characteristics And Sensitivity Trades For NASA's Mars Design Reference Architecture (DRA) 5.0 Study

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.; McCurdy, David R.; Packard, Thomas W.

2009-01-01

This paper summarizes Phase I and II analysis results from NASA's recent Mars DRA 5.0 study which re-examined mission, payload and transportation system requirements for a human Mars landing mission in the post-2030 timeframe. Nuclear thermal rocket (NTR) propulsion was again identified as the preferred in-space transportation system over chemical/aerobrake because of its higher specific impulse (I(sub sp)) capability, increased tolerance to payload mass growth and architecture changes, and lower total initial mass in low Earth orbit (IMLEO) which is important for reducing the number of Ares-V heavy lift launches and overall mission cost. DRA 5.0 features a long surface stay (approximately 500 days) split mission using separate cargo and crewed Mars transfer vehicles (MTVs). All vehicles utilize a common core propulsion stage with three 25 klbf composite fuel NERVA-derived NTR engines (T(sub ex) approximately 2650 - 2700 K, p(sub ch) approximately 1000 psia, epsilon approximately 300:1, I(sub sp) approximately 900 - 910 s, engine thrust-toweight ratio approximately 3.43) to perform all primary mission maneuvers. Two cargo flights, utilizing 1-way minimum energy trajectories, pre-deploy a cargo lander to the surface and a habitat lander into a 24-hour elliptical Mars parking orbit where it remains until the arrival of the crewed MTV during the next mission opportunity (approximately 26 months later). The cargo payload elements aerocapture (AC) into Mars orbit and are enclosed within a large triconicshaped aeroshell which functions as payload shroud during launch, then as an aerobrake and thermal protection system during Mars orbit capture and subsequent entry, descent and landing (EDL) on Mars. The all propulsive crewed MTV is a 0-gE vehicle design that utilizes a fast conjunction trajectory that allows approximately 6-7 month 1-way transit times to and from Mars. Four 12.5 kW(sub e) per 125 square meter rectangular photovoltaic arrays provide the crewed MTV with approximately 50 kW(sub e) of electrical power in Mars orbit for crew life support and spacecraft subsystem needs. Vehicle assembly involves autonomous Earth orbit rendezvous and docking between the propulsion stages, in-line propellant tanks and payload elements. Nine Ares-V launches -- five for the two cargo MTVs and four for the crewed MTV -- deliver the key components for the three MTVs. Details on mission, payload, engine and vehicle characteristics and requirements are presented and the results of key trade studies are discussed.

Concept for a Lunar Transfer Vehicle for Small Satellite Delivery to the Moon from the International Space Station

NASA Technical Reports Server (NTRS)

Elliott, John; Alkalai, Leon

2010-01-01

The International Space Station (ISS) has developed as a very capable center for scientific research in Lower Earth Orbit. An additional potential of the ISS that has not thus far been exploited, is the use of this orbiting plat-form for the assembly and launching of vehicles that could be sent to more distant destinations. This paper reports the results of a recent study that looked at an architecture and conceptual flight system design for a lunar transfer vehicle (LTV) that could be delivered to the ISS in segments, assembled, loaded with payload and launched from the ISS with the objective of delivering multiple small and micro satellites to lunar orbit. The design of the LTV was optimized for low cost and high payload capability, as well as ease of assembly. The resulting design would use solar electric propulsion (SEP) to carry a total payload mass of 250 kg from the ISS to a 100 km lunar orbit. A preliminary concept of operations was developed considering currently available delivery options and ISS capabili-ties that should prove flexible enough to accommodate a variety of payloads and missions. This paper will present an overview of the study, including key trades, mission and flight system design, and notional operational concept.

Mass drivers. 3: Engineering

NASA Technical Reports Server (NTRS)

Arnold, W.; Bowen, S.; Cohen, S.; Fine, K.; Kaplan, D.; Kolm, M.; Kolm, H.; Newman, J.; Oneill, G. K.; Snow, W.

1979-01-01

The last of a series of three papers by the Mass-Driver Group of the 1977 Ames Summer Study is presented. It develops the engineering principles required to implement the basic mass-driver. Optimum component mass trade-offs are derived from a set of four input parameters, and the program used to design a lunar launcher. The mass optimization procedures is then incorporated into a more comprehensive mission optimization program called OPT-4, which evaluates an optimized mass-driver reaction engine and its performance in a range of specified missions. Finally, this paper discusses, to the extent that time permitted, certain peripheral problems: heating effects in buckets due to magnetic field ripple; an approximate derivation of guide force profiles; the mechanics of inserting and releasing payloads; the reaction mass orbits; and a proposed research and development plan for implementing mass drivers.

Mass Property Measurements of the Mars Science Laboratory Rover

NASA Technical Reports Server (NTRS)

Fields, Keith

2012-01-01

The NASA/JPL Mars Science Laboratory (MSL) spacecraft mass properties were measured on a spin balance table prior to launch. This paper discusses the requirements and issues encountered with the setup, qualification, and testing using the spin balance table, and the idiosyncrasies encountered with the test system. The final mass measurements were made in the Payload Hazardous Servicing Facility (PHSF) at Kennedy Space Center on the fully assembled and fueled spacecraft. This set of environmental tests required that the control system for the spin balance machine be at a remote location, which posed additional challenges to the operation of the machine

Evidence for a maximum mass cut-off in the neutron star mass distribution and constraints on the equation of state

NASA Astrophysics Data System (ADS)

Alsing, Justin; Silva, Hector O.; Berti, Emanuele

2018-04-01

We infer the mass distribution of neutron stars in binary systems using a flexible Gaussian mixture model and use Bayesian model selection to explore evidence for multi-modality and a sharp cut-off in the mass distribution. We find overwhelming evidence for a bimodal distribution, in agreement with previous literature, and report for the first time positive evidence for a sharp cut-off at a maximum neutron star mass. We measure the maximum mass to be 2.0M⊙ < mmax < 2.2M⊙ (68%), 2.0M⊙ < mmax < 2.6M⊙ (90%), and evidence for a cut-off is robust against the choice of model for the mass distribution and to removing the most extreme (highest mass) neutron stars from the dataset. If this sharp cut-off is interpreted as the maximum stable neutron star mass allowed by the equation of state of dense matter, our measurement puts constraints on the equation of state. For a set of realistic equations of state that support >2M⊙ neutron stars, our inference of mmax is able to distinguish between models at odds ratios of up to 12: 1, whilst under a flexible piecewise polytropic equation of state model our maximum mass measurement improves constraints on the pressure at 3 - 7 × the nuclear saturation density by ˜30 - 50% compared to simply requiring mmax > 2M⊙. We obtain a lower bound on the maximum sound speed attained inside the neutron star of c_s^max > 0.63c (99.8%), ruling out c_s^max < c/√{3} at high significance. Our constraints on the maximum neutron star mass strengthen the case for neutron star-neutron star mergers as the primary source of short gamma-ray bursts.

Emerging Array Antenna Technologies at JPL

NASA Technical Reports Server (NTRS)

Huang, J.

1998-01-01

JPL/NASA's Earth remote sensing and deep-space exploration programs have been placing emphasis on their spacecraft's high-gain and large-aperture antennas. At the same time, however, low mass and small storage volume are demanded in order to reduce payload weight and reduce shroud size and thus reduce launch cost.

Game Changing: NASA's Space Launch System and Science Mission Design

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2013-01-01

NASA s Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as "monolithic" telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

Game changing: NASA's space launch system and science mission design

NASA Astrophysics Data System (ADS)

Creech, S. D.

NASA's Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher characteristic energy (C3) energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as “ monolithic” telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

A broad-band microseismometer for planetary operations

NASA Technical Reports Server (NTRS)

Banerdt, W. B.; Vanzandt, T.; Kaiser, W. J.; Kenny, T. W.

1993-01-01

There has recently been renewed interest in the development of instrumentation for making measurements on the surface of Mars. This is due to the Mars Environmental Survey (MESUR) Mission, for which approximately 16 small, long-lived (2-10 years), relatively inexpensive surface stations will be deployed in a planet-wide network. This will allow the investigation of processes (such as seismology and meteorology) which require the simultaneous measurement of phenomena at many widely spaced locations on the surface over a considerable length of time. Due to the large number of vehicles involved, the mass, power, and cost of the payload will be severely constrained. A seismometer has been identified as one of the highest priority instruments in the MESUR straw-man payload. The requirements for an effective seismic experiment on Mars place a number of constraints on any viable sensor design. First, a large number of sensors must be deployed in a long-lived global network in order to be able to locate many events reliably, provide good spatial sampling of the interior, and increase the probability of seismic detection in the event of localized seismicity and/or high attenuation. From a practical standpoint, this means that individual surface stations will necessarily be constrained in terms of cost, mass, and power. Landing and thermal control systems will probably be simple, in order to minimize cost, resulting in large impact accelerations and wide daily and seasonal thermal swings. The level of seismic noise will determine the maximum usable sensitivity for seismometer. Unfortunately, the ambient seismic noise level for Mars is not well known. However lunar seismic noise levels are several orders of magnitude below that of the Earth. Sensitivities on the order of 10(exp -11)g over a bandwidth of .04 to 20 Hz are thought to be necessary to fulfill the science objectives for a seimometer placed on the Martian surface. Silicon micromachined sensor technology offers techniques for the fabrication of monolithic, robust, compact, lower power and mass accelerometers. Conventional micro-machined accelerometers have been developed and are commercially available for high frequency and large acceleration measurements. The new seismometer we are developing incorporates certain principles of conventional silicon micromachined accelerometer technology. However, currently available silicon micromachined sensors offer inadequate sensitivity and bandwidth for the Mars seismometer application. Our implementation of an advanced silicon micromachined seismometer is based on principles recently developed at JPL for high-sensitivity position sensor technology.

Antares Rocket Test Launch

NASA Image and Video Library

2013-04-21

NASA Administrator Charles Bolden and NASA Deputy Administrator Lori Garver and other guests react after having watched the successful launch of the Orbital Sciences Corporation Antares rocket from the Mid-Atlantic Regional Spaceport (MARS) at the NASA Wallops Flight Facility in Virginia, Sunday, April 21, 2013. The test launch marked the first flight of Antares and the first rocket launch from Pad-0A. The Antares rocket delivered the equivalent mass of a spacecraft, a so-called mass simulated payload, into Earth's orbit. Photo Credit: (NASA/Bill Ingalls)

Control System Upgrade for a Mass Property Measurement Facility

NASA Technical Reports Server (NTRS)

Chambers, William; Hinkle, R. Kenneth (Technical Monitor)

2002-01-01

The Mass Property Measurement Facility (MPMF) at the Goddard Space Flight Center has undergone modifications to ensure the safety of Flight Payloads and the measurement facility. The MPMF has been technically updated to improve reliability and increase the accuracy of the measurements. Modifications include the replacement of outdated electronics with a computer based software control system, the addition of a secondary gas supply in case of a catastrophic failure to the gas supply and a motor controlled emergency stopping feature instead of a hard stop.

NASA's Space Launch System: An Evolving Capability for Exploration

NASA Technical Reports Server (NTRS)

Robinson, Kimberly F.; Hefner, Keith; Hitt, David

2015-01-01

Designed to enable human space exploration missions, including eventually landings on Mars, NASA's Space Launch System (SLS) represents a unique launch capability with a wide range of utilization opportunities, from delivering habitation systems into the "proving ground" of lunar-vicinity space to enabling high-energy transits through the outer solar system. Substantial progress has been made toward the first launch of the initial configuration of SLS, which will be able to deliver more than 70 metric tons of payload into low Earth orbit (LEO). Preparations are also underway to evolve the vehicle into more powerful configurations, culminating with the capability to deliver more than 130 metric tons to LEO. Even the initial configuration of SLS will be able to deliver greater mass to orbit than any contemporary launch vehicle, and the evolved configuration will have greater performance than the Saturn V rocket that enabled human landings on the moon. SLS will also be able to carry larger payload fairings than any contemporary launch vehicle, and will offer opportunities for co-manifested and secondary payloads. Because of its substantial mass-lift capability, SLS will also offer unrivaled departure energy, enabling mission profiles currently not possible. The basic capabilities of SLS have been driven by studies on the requirements of human deep-space exploration missions, and continue to be validated by maturing analysis of Mars mission options, including the Global Exploration Roadmap. Early collaboration with science teams planning future decadal-class missions have contributed to a greater understanding of the vehicle's potential range of utilization. As SLS draws closer to its first launch, the Program is maturing concepts for future capability upgrades, which could begin being available within a decade. These upgrades, from multiple unique payload accommodations to an upper stage providing more power for inspace propulsion, have ramifications for a variety of missions, from human exploration to robotic science.

Surface Experiments on a Jupiter Trojan Asteroid in the Solar Powered Sail Mission

NASA Astrophysics Data System (ADS)

Okada, Tatsuaki

2016-04-01

Introduction: A new mission to a Jupiter Trojan asteroid is under study us-ing a solar-powered sail (SPS), and a science lander is being investigated in the joint study between Japan and Europe [1]. We present here the key sci-entific objectives and the strawman payloads of science experiments on the asteroid. Science Objectives: Jupiter Trojan asteroids are located around the Sun-Jupiter Lagrange points (L4 or L5) and most of them are classified as D- or P-type in asteroid taxonomy, but their origin still remains unknown. A classi-cal (static) model of solar system evolution indicates that they were formed around the Jupiter region and survived until now as the outer end members of asteroids. A new (dynamical) model such as Nice model suggests that they were formed at the far end of the solar system and transferred inward due to dynamical migration of giant planets [2]. Therefore physical, miner-alogical, and isotopic studies of surface materials and volatile compounds could solve their origin, and then the solar system formation [3]. Strawman Payloads: The SPS orbiter will be able to carry a 100 kg class lander with 20 kg mission payloads. Just after landing of the lander, geolog-ical, mineralogical, and geophysical observations will be performed to char-acterize the site using a panoramic optical camera, an infrared hyperspectral imager, a magnetometer, and a thermal radiometer. The surface and subsur-face materials of the asteroid will be collected into a carousel by the bullet-type and the pneumatic drill type samplers, respectively. Samples in the carousel will be investigated by a visible and an infrared microscope, and transferred for performing high resolution mass spectrometry (HRMS). Mass resolution m/dm > 30,000 is expected to investigate isotopic ratios of D/H, 15N/14N, and 18O/16O, as well as molecules from organic matters. A set of strawman payloads are tentatively determined during the lander system study [4]. The constraints to select the strawman payloads have the total mass of 20 kg, and the total consumption energy of 600 WHr. In the SPS mission, the sample-return is also studied as an option, and the lander should bring the mechanisms for sample collection and sample transfer to the mother ship. [1] Mori O. et al. (2015) 11th Low-Cost Planetary Missions Conf., S3-10. [2] Morbidelli A. et al. (2005) Nature 435, 462-466. [3] Yano H. et al., (2014) CO-SPAR 2014, B0.4-2-14. [4] Mori O. et al., Lunar Planet. Sci. Conf., 47, #1822.

High performance optical payloads for microsatellites

NASA Astrophysics Data System (ADS)

Geyl, Roland; Rodolfo, Jacques; Girault, Jean-Philippe

2017-09-01

Safran is presenting two concepts of optical payloads for microsatellites combining high performances and extremely compact volume. The first one offer 10-m Ground Sampling Distance (GSD) over 60x40 km2 area from 600 km orbit optimized for twilight conditions. The second one is offering a much higher resolution of 1.8-m over 11x7,5 km2 area from the same 600 km orbit. The two concepts are based on advanced innovative diffraction limited optical system packaged in a unique very compact volume lower than 8U = 200x200x200 mm making them the ideal solution for 15- 100 kg microsatellites. The maximum number of pixels is served to the end-user space imagery community thanks to 35 mm Full Frame sensors offering, as of today, 6000x4000 pixels. Up to 10 spectral bands from 475 to 900 nm can be offered thanks to 2D structured filters.

NASA standard GAS Can satellite. [Get-Away Special canister for STS Orbiter

NASA Technical Reports Server (NTRS)

Cudmore, Patrick H.; Mcintosh, W.; Edison, M.; Nichols, S.; Mercier, E.

1989-01-01

The Get-Away Special canister (GAS Can) satellite is a small, (150 lb) low-cost satellite making it possible for commercial and scientific institutions to conduct experiments in space on an economical and short-term basis. The current model is called Xsat (Exceptional Satellite) and is designed to be launched from a GAS canister on the STS Orbiter; also provided is a low-cost automated PC-operated ground station for commercial, scientific, and government users. The Xsat structure is diagrammed, and details such as payload interface, weight restrictions, and structural loads are described in detail, pointing out that Xsat has a maximum payload weight of 50 lbs, and has a natural vibration frequency of around 45 Hz, with a minimum requiremet of 35 Hz. Thermal designs, power system, electronics, computer design and bus system, and satellite operations are all outlined.

KSC-00pp1682

NASA Image and Video Library

2000-11-10

In the Space Station Processing Facility, an overhead crane lifts the P6 integrated truss segment from a workstand to place it in the payload transport canister for transfer to Launch Pad 39B. There it will be placed in Endeavour’s payload bay for launch on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1683

NASA Image and Video Library

2000-11-10

In the Space Station Processing Facility, an overhead crane moves the P6 integrated truss segment to a payload transport canister for transfer to Launch Pad 39B. There it will be placed in Endeavour’s payload bay for launch on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1738

NASA Image and Video Library

2000-11-14

The doors of the payload transport canister are open wide in the payload changeout room on Launch Pad 39B. Revealed is the P6 integrated truss segment, which will fly on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch of STS-97 is scheduled for Nov. 30 at 10:06 p.m. EST

Building Aerodynamic Databases for the SLS Design Process

NASA Technical Reports Server (NTRS)

Rogers, Stuart; Dalle, Derek J.; Lee, Henry; Meeroff, Jamie; Onufer, Jeffrey; Chan, William; Pulliam, Thomas

2017-01-01

NASA's new Space Launch System (SLS) will be the first rocket since the Saturn V (1967-1973) to carry astronauts beyond low earth orbit-and will carry 10% more payload than Saturn V and three times the payload of the space shuttle. The SLS configuration consists of a center core and two solid rocket boosters that separate from the core as their fuel is exhausted two minutes after lift-off. During these first two minutes of flight, the vehicle powers its way through strong shock waves as it accelerates past the speed of sound, then pushes beyond strong aerodynamic loads at the maximum dynamic pressure, and is ultimately enveloped by gaseous plumes from the booster-separation motors. The SLS program relies on computational fluid dynamic (CFD) simulations to provide much of the data needed to build aerodynamic databases describing the structural load distribution, surface pressures, and aerodynamic forces on the vehicle.

Numerical Simulation of Parachute Inflation Process by IB Method

NASA Astrophysics Data System (ADS)

Miyoshi, Masaya; Mori, Koichi; Nakamura, Yoshiaki

In the present study the deformation and motion of a parachute in the process of inflation are simulated by applying the immersed boundary technique in a fluid-structure coupling solver. It was found from simulated results that the canopy is first inflated in the normal direction to the uniform flow (in the lateral direction), and then its apex is pulled by a vortex ring generated near the canopy's outer surface due to its negative pressure. After the end of this inflation process, the canopy moves in the tangential direction to the spherical surface, the center of which is located at the payload location. This motion is caused by the breakup of an initial axisymmetric vortex, where many vortices are generated from the shear layer. The predicted maximum parachute opening force is twice as large as the payload force in the steady state, which is in good agreement with experiment.

Using Technology Readiness Level (TRL), Life Cycle Cost (LCC), and Other Metrics to Supplement Equivalent System Mass (ESM) in Advanced Life Support (ALS)

NASA Technical Reports Server (NTRS)

Jones, Harry

2003-01-01

The ALS project plan goals are reducing cost, improving performance, and achieving flight readiness. ALS selects projects to advance the mission readiness of low cost, high performance technologies. The role of metrics is to help select good projects and report progress. The Equivalent Mass (EM) of a system is the sum of the estimated mass of the hardware, of its required materials and spares, and of the pressurized volume, power supply, and cooling system needed to support the hardware in space. EM is the total payload launch mass needed to provide and support a system. EM is directly proportional to the launch cost.

On-Line, Gyro-Based, Mass-Property Identification for Thruster-Controlled Spacecraft Using Recursive Least Squares

NASA Technical Reports Server (NTRS)

Wilson, Edward; Lages, Chris; Mah, Robert; Clancy, Daniel (Technical Monitor)

2002-01-01

Spacecraft control, state estimation, and fault-detection-and-isolation systems are affected by unknown v aerations in the vehicle mass properties. It is often difficult to accurately measure inertia terms on the ground, and mass properties can change on-orbit as fuel is expended, the configuration changes, or payloads are added or removed. Recursive least squares -based algorithms that use gyro signals to identify the center of mass and inverse inertia matrix are presented. They are applied in simulation to 3 thruster-controlled vehicles: the X-38 and Mini-AERCam under development at NASA-JSC, and the SAM, an air-bearing spacecraft simulator at the NASA-Ames Smart Systems Research Lab (SSRL).

Robotic planetary science missions enabled with small NTR engine/stage technologies

NASA Astrophysics Data System (ADS)

Borowski, Stanley K.

1995-10-01

The high specific impulse (Isp) and engine thrust-to-weight ratio of liquid hydrogen (LH2)-cooled nuclear thermal rocket (NTR) engines makes them ideal for upper stage applications to difficult robotic planetary science missions. A small 15 thousand pound force (klbf) NTR engine using a uranium-zirconium-niobium 'ternary carbide' fuel (Isp approximately 960 seconds at approximately 3025K) developed in the Commonwealth of Independent States (CIS) is examined and its use on an expendable injection stage is shown to provide major increases in payload delivered to the outer planets (Saturn, Uranus, Neptune and Pluto). Using a single 'Titan IV-class' launch vehicle, with a lift capability to low Earth orbit (LEO) of approximately 20 metric tons (t), an expendable NTR upper stage can inject two Pluto 'Fast Flyby' spacecraft (PFF/SC) plus support equipment-combined mass of approximately 508 kg--on high energy, '6.5-9.2 year' direct trajectory missions to Pluto. A conventional chemical propulsion mission would use a liquid oxygen (LOX)/LH2 'Centaur' upper stage and two solid rocket 'kick motors' to inject a single PFF/SC on the same Titan IV launch vehicle. For follow on Pluto missions, the NTR injection stage would utilize a Jupiter 'gravity assist' (JGA) maneuver to launch a LOX/liquid methane (CH4) capture stage (Isp approximately 375 seconds) and a Pluto 'orbiter' spacecraft weighing between approximately 167-312 kg. With chemical propulsion, a Pluto orbiter mission is not a viable option because c inadequate delivered mass. Using a 'standardized' NTR injection stage and the same single Titan IV launch scenario, 'direct flight' (no gravity assist) orbiter missions to Saturn, Uranus and Neptune are also enabled with transit times of 2.3, 6.6, and 12.6 years, respectively. Injected mass includes a storable, nitrogen tetroxide/monomethyl hydrazine (N2O4/MMH) capture stage (Isp approximately 330 seconds) and orbiter payloads 340 to 820% larger than that achievable using a LOX/LH2-fueled injection stage. The paper discusses NTR technology and mission characteristics, shows NTR stage and payload accommodations within the 26.2 m long Titan IV payload fairing, and discusses NTR stage performance as a function of assumed cryogenic tank technology.

Robotic Planetary Science Missions Enabled with Small NTR Engine/Stage Technologies

NASA Technical Reports Server (NTRS)

Borowski, Stanley K.

1995-01-01

The high specific impulse (Isp) and engine thrust-to-weight ratio of liquid hydrogen (LH2)-cooled nuclear thermal rocket (NTR) engines makes them ideal for upper stage applications to difficult robotic planetary science missions. A small 15 thousand pound force (klbf) NTR engine using a uranium-zirconium-niobium 'ternary carbide' fuel (Isp approximately 960 seconds at approximately 3025K) developed in the Commonwealth of Independent States (CIS) is examined and its use on an expendable injection stage is shown to provide major increases in payload delivered to the outer planets (Saturn, Uranus, Neptune and Pluto). Using a single 'Titan IV-class' launch vehicle, with a lift capability to low Earth orbit (LEO) of approximately 20 metric tons (t), an expendable NTR upper stage can inject two Pluto 'Fast Flyby' spacecraft (PFF/SC) plus support equipment-combined mass of approximately 508 kg--on high energy, '6.5-9.2 year' direct trajectory missions to Pluto. A conventional chemical propulsion mission would use a liquid oxygen (LOX)/LH2 'Centaur' upper stage and two solid rocket 'kick motors' to inject a single PFF/SC on the same Titan IV launch vehicle. For follow on Pluto missions, the NTR injection stage would utilize a Jupiter 'gravity assist' (JGA) maneuver to launch a LOX/liquid methane (CH4) capture stage (Isp approximately 375 seconds) and a Pluto 'orbiter' spacecraft weighing between approximately 167-312 kg. With chemical propulsion, a Pluto orbiter mission is not a viable option because c inadequate delivered mass. Using a 'standardized' NTR injection stage and the same single Titan IV launch scenario, 'direct flight' (no gravity assist) orbiter missions to Saturn, Uranus and Neptune are also enabled with transit times of 2.3, 6.6, and 12.6 years, respectively. Injected mass includes a storable, nitrogen tetroxide/monomethyl hydrazine (N2O4/MMH) capture stage (Isp approximately 330 seconds) and orbiter payloads 340 to 820% larger than that achievable using a LOX/LH2-fueled injection stage. The paper discusses NTR technology and mission characteristics, shows NTR stage and payload accommodations within the 26.2 m long Titan IV payload fairing, and discusses NTR stage performance as a function of assumed cryogenic tank technology.

Historical Reveiw of Interstellar Probe Concepts and Examination of Payload Mass Considerations for Different System Architectures

NASA Astrophysics Data System (ADS)

Long, K.

2017-12-01

The ability to send a space probe beyond the Voyager probes, through the interstellar medium and towardsthe distant stars, has long been the ambition of both the science ction literature but also a small community ofadvocates that have argued for a broader and deeper vision of space exploration that goes outside of our SolarSystem. In this paper we discuss some of the historical interstellar probe concepts which are propelled usingdierent types of propulsion technology, from energetic reaction engines to directed energy beaming, and considerthe payload mass associated with such concepts. We compare and contrast the dierent design concepts, payloadmass fractions, powers and energies and discuss the implications for robotic space exploration within the stellarneighbourhood. Finally, we consider the Breakthrough Starshot initiative, which proposes to send a Gram-scalelaser driven spacecraft to the Alpha Centauri system in a 20 year mission travelling at v 0.2c. We show howthis is a good start in pushing our robotic probes towards interstellar destinations, but also discuss the potentialfor scaling up this systems architecture to missions closer at home, or higher mass missions wider aeld. This is apresentation for the American Geophysical Union at the AGU Fall meeting, New Orleans, 11-15 December 2017,Special Session on the Interstellar Probe Missions.Keywords: Interstellar Probe, Breakthrough Starshot

The C3PO project: a laser communication system concept for small satellites

NASA Astrophysics Data System (ADS)

d'Humières, Benoît; Esmiller, Bruno; Gouy, Yann; Steck, Emilie; Quintana, Crisanto; Faulkner, Graham; O'Brien, Dominic; Sproll, Fabian; Wagner, Paul; Hampf, Daniel; Riede, Wolfgang; Salter, Michael; Wang, Qin; Platt, Duncan; Jakonis, Darius; Piao, Xiaoyu; Karlsson, Mikael; Oberg, Olof; Petermann, Ingemar; Michalkiewicz, Aneta; Krezel, Jerzy; Debowska, Anna; Thueux, Yoann

2017-02-01

The satellite market is shifting towards smaller (micro and nanosatellites), lowered mass and increased performance platforms. Nanosatellites and picosatellites have been used for a number of new, innovative and unique payloads and missions. This trend requires new concepts for a reduced size, a better performance/weight ratio and a reduction of onboard power consumption. In this context, disruptive technologies, such as laser-optical communication systems, are opening new possibilities. This paper presents the C3PO1 system, "advanced Concept for laser uplink/ downlink CommuniCation with sPace Objects", and the first results of the development of its key technologies. This project targets the design of a communications system that uses a ground-based laser to illuminate a satellite, and a Modulating Retro-Reflector (MRR) to return a beam of light modulated by data to the ground. This enables a downlink, without a laser source on the satellite. This architecture suits well to small satellite applications so as high data rates are potentially provided with very low board mass. C3PO project aims to achieve data rates of 1Gbit/s between LEO satellites and Earth with a communication payload mass of less than 1kilogram. In this paper, results of the initial experiments and demonstration of the key technologies will be shown.

Mars Rotorcraft: Possibilities, Limitations, and Implications For Human/Robotic Exploration

NASA Technical Reports Server (NTRS)

Young, Larry A.; Aiken, Edwin; Lee, Pascal; Briggs, Geoffrey

2005-01-01

Several research investigations have examined the challenges and opportunities in the use of small robotic rotorcraft for the exploration of Mars. To date, only vehicles smaller than 150 kg have been studied. This paper proposes to examine the question of maximum Mars rotorcraft size, range, and payload/cargo capacity. Implications for the issue of whether or not (from an extreme design standpoint) a manned Mars rotorcraft is viable are also discussed.

Minuteman 2 launched small satellite

NASA Technical Reports Server (NTRS)

Chan, Sunny; Hinders, Kriss; Martin, Trent; Mcmillian, Shandy; Sharp, Brad; Vajdos, Greg

1994-01-01

The goal of LEOSat Industries' Spring 1994 project was to design a small satellite that has a strong technology demonstration or scientific justification and incorporates a high level of student involvement. The satellite is to be launched into low earth orbit by the converted Minuteman 2 satellite launcher designed by Minotaur Designs, Inc. in 1993. The launch vehicle shroud was modified to a height of 90 inches, a diameter of 48 inches at the bottom and 35 inches at the top for a total volume of 85 cubic feet. The maximum allowable mass of the payload is about 1100 lb., depending on the launch site, orbit altitude, and inclination. The satellite designed by LEOSat Industries is TerraSat, a remote-sensing satellite that will provide information for use in space-based earth studies. It will consist of infrared and ultraviolet/visible sensors similar to the SDI-developed sensors being tested on Clementine. The sensors will be mounted on the Defense Systems, Inc. Standard Satellite-1 spacecraft bus. LEOSat has planned for two satellites orbiting the Earth with trajectories similar to that of LANDSAT 5. The semi-major axis is 7080 kilometers, the eccentricity is 0, and the inclination is 98.2 degrees. The estimated mass of TerraSat is 145 kilograms and the estimated volume is 1.8 cubic meters. The estimated cost of TerraSat is $13.7 million. The projected length of time from assembly of the sensors to launch of the spacecraft is 13 months.

NASA's Space Launch System: Deep-Space Opportunities for SmallSats

NASA Technical Reports Server (NTRS)

Robinson, Kimberly F.; Schorr, Andrew A.

2017-01-01

Designed for human exploration missions into deep space, NASA's Space Launch System (SLS) represents a new spaceflight infrastructure asset, enabling a wide variety of unique utilization opportunities. While primarily focused on launching the large systems needed for crewed spaceflight beyond Earth orbit, SLS also offers a game-changing capability for the deployment of small satellites to deep-space destinations, beginning with its first flight. Currently, SLS is making rapid progress toward readiness for its first launch in two years, using the initial configuration of the vehicle, which is capable of delivering 70 metric tons (t) to Low Earth Orbit (LEO). On its first flight test of the Orion spacecraft around the moon, accompanying Orion on SLS will be small-satellite secondary payloads, which will deploy in cislunar space. The deployment berths are sized for "6U" CubeSats, and on EM-1 the spacecraft will be deployed into cislunar space following Orion separate from the SLS Interim Cryogenic Propulsion Stage. Payloads in 6U class will be limited to 14 kg maximum mass. Secondary payloads on EM-1 will be launched in the Orion Stage Adapter (OSA). Payload dispensers will be mounted on specially designed brackets, each attached to the interior wall of the OSA. For the EM-1 mission, a total of fourteen brackets will be installed, allowing for thirteen payload locations. The final location will be used for mounting an avionics unit, which will include a battery and sequencer for executing the mission deployment sequence. Following the launch of EM-1, deployments of the secondary payloads will commence after sufficient separation of the Orion spacecraft to the upper stage vehicle to minimize any possible contact of the deployed cubesats to Orion. Currently this is estimated to require approximately 4 hours. The allowed deployment window for the cubesats will be from the time the upper stage disposal maneuvers are complete to up to 10 days after launch. The upper stage will fly past the moon at a perigee of approximately 100km, and this closest approach will occur about 5 days after launch. The limiting factor for the latest deployment time is the available power in the sequencer system. Several NASA Mission Directorates were involved in the development of programs for the competition, selection, and development of EM-1 payloads that support directorate priorities. CubeSat payloads on EM-1 will include both NASA research experiments and spacecraft developed by industry, international and potentially academia partners. The Human Exploration and Operations Mission Directorate (HEOMD) Advanced Exploration Systems (AES) Division was allocated five payload opportunities on the EM-1 mission. Near Earth Asteroid (NEA) Scout is designed to rendezvous with and characterize a candidate NEA. A solar sail, an innovation the spacecraft will demonstrated for the CubeSat class, will provide propulsion. Lunar Flashlight will use a green propellant system and will search for potential ice deposits in the moon's permanently shadowed craters. BioSentinel is a yeast radiation biosensor, planned to measure the effects of space radiation on deoxyribonucleic acid (DNA). Lunar Icecube, a collaboration with Morehead State University, will prospect for water in ice, liquid, and vapor forms as well as other lunar volatiles from a low-perigee, highly inclined lunar orbit using a compact Infrared spectrometer. Skyfire, a partnership with Lockheed Martin, is a technology demonstration mission that will perform a lunar flyby, collecting spectroscopy, and thermography data to address questions related to surface characterization, remote sensing, and site selection. NASA's Space Technology Mission Directorate (STMD) was allocated three payload opportunities on the EM-1 mission. These slots will be filled via the 2 Centennial Challenges Program, NASA's flagship program for technology prize competitions, which directly engages the public, academia, and industry in open prize competitions to stimulate innovation. The NASA Science Mission Directorate (SMD) was allocated two payload opportunities on the EM-1 mission. The CubeSat Mission to Study Solar Particles (CuSP) payload will study the sources and acceleration mechanisms of solar and interplanetary particles in near-Earth orbit, support space weather research by determining proton radiation levels during Solar Energetic Particle (SEP) events and identifying suprathermal properties that could help predict geomagnetic storms. The LunaH-Map payload will help scientists understand the quantity of H-bearing materials in lunar cold traps (10 km), determine the concentration of H-bearing materials with 1m depth, and constrain the vertical distribution of H-bearing materials. The final three payload opportunities for the EM-1 mission were allocated for NASA's international space agency counterparts. The flight opportunities are intended to benefit the international space agency and NASA as well as further the collective space exploration goals. ArgoMoon is sponsored by ESA/ASI and will fly along with the ICPS on its disposal trajectory to perform proximity operations with the ICPS post-disposal, take external imagery of engineering and historical significance, and perform an optical communications demonstration. EQUULEUS, sponsored by JAXA, will fly to a libration orbit around the Earth-Moon L2 point and demonstrate trajectory control techniques within the Sun-Earth-Moon region for the first time by a nano spacecraft. The mission will also contribute to the future human exploration scenario by understanding the radiation environment in geospace and deep space, characterizing the flux of impacting meteors on the far side of the moon, and demonstrating the future deep space exploration scenario using the "deep space port" at Lagrange points. OMOTENASHI, also sponsored by JAXA, will land the smallest lunar lander to date on the lunar surface to demonstrate the feasibility of the hardware for distributed cooperative exploration system. Small landers will enable multi-point exploration, which is complimentary with large-scale human exploration. Once on the lunar surface, the OMOTENASHI spacecraft will observe the radiation and soil environments of the lunar surface by active radiation measurements and soil shear measurements. Following EM-1, Space Launch System will evolve to the more-powerful Block 1B configuration, which uses a new Exploration Upper Stage to increase the vehicle's LEO payload capability from 70 t to 105 t. With that transition, the Orion Stage Adapter, which will carry the secondary payloads on EM-1, will be phased out, and a new Universal Stage Adapter will be introduced, creating opportunities for flying larger secondary payloads. This paper will provide a brief status of SLS progress toward first launch; an overview of smallsat accommodations, integration, and operations on EM-1; information about the specific payloads flying on that launch; and a discussion of future accommodations and opportunities for secondary payloads on SLS for Exploration Mission-2 and beyond.

NASA's Space Launch System: Deep-Space Delivery for Smallsats

NASA Technical Reports Server (NTRS)

Robinson, Kimberly F.; Norris, George

2017-01-01

Designed for human exploration missions into deep space, NASA's Space Launch System (SLS) represents a new spaceflight infrastructure asset, enabling a wide variety of unique utilization opportunities. While primarily focused on launching the large systems needed for crewed spaceflight beyond Earth orbit, SLS also offers a game-changing capability for the deployment of small satellites to deep-space destinations, beginning with its first flight. Currently, SLS is making rapid progress toward readiness for its first launch in two years, using the initial configuration of the vehicle, which is capable of delivering 70 metric tons (t) to Low Earth Orbit (LEO). On its first flight test of the Orion spacecraft around the moon, accompanying Orion on SLS will be small-satellite secondary payloads, which will deploy in cislunar space. The deployment berths are sized for "6U" CubeSats, and on EM-1 the spacecraft will be deployed into cislunar space following Orion separate from the SLS Interim Cryogenic Propulsion Stage. Payloads in 6U class will be limited to 14 kg maximum mass. Secondary payloads on EM-1 will be launched in the Orion Stage Adapter (OSA). Payload dispensers will be mounted on specially designed brackets, each attached to the interior wall of the OSA. For the EM-1 mission, a total of fourteen brackets will be installed, allowing for thirteen payload locations. The final location will be used for mounting an avionics unit, which will include a battery and sequencer for executing the mission deployment sequence. Following the launch of EM-1, deployments of the secondary payloads will commence after sufficient separation of the Orion spacecraft to the upper stage vehicle to minimize any possible contact of the deployed CubeSats to Orion. Currently this is estimated to require approximately 4 hours. The allowed deployment window for the CubeSats will be from the time the upper stage disposal maneuvers are complete to up to 10 days after launch. The upper stage will fly past the moon at a perigee of approximately 100km, and this closest approach will occur about 5 days after launch. The limiting factor for the latest deployment time is the available power in the sequencer system. Several NASA Mission Directorates were involved in the development of programs for the competition, selection, and development of EM-1 payloads that support directorate priorities. CubeSat payloads on EM-1 will include both NASA research experiments and spacecraft developed by industry, international and potentially academia partners. The Human Exploration and Operations Mission Directorate (HEOMD) Advanced Exploration Systems (AES) Division was allocated five payload opportunities on the EM-1 mission. Near Earth Asteroid (NEA) Scout is designed to rendezvous with and characterize a candidate NEA. A solar sail, an innovation the spacecraft will demonstrated for the CubeSat class, will provide propulsion. Lunar Flashlight will use a green propellant system and will search for potential ice deposits in the moon's permanently shadowed craters. BioSentinel is a yeast radiation biosensor, planned to measure the effects of space radiation on deoxyribonucleic acid (DNA). Lunar Icecube, a collaboration with Morehead State University, will prospect for water in ice, liquid, and vapor forms as well as other lunar volatiles from a low-perigee, highly inclined lunar orbit using a compact Infrared spectrometer. Skyfire, a partnership with Lockheed Martin, is a technology demonstration mission that will perform a lunar flyby, collecting spectroscopy, and thermography data to address questions related to surface characterization, remote sensing, and site selection. NASA's Space Technology Mission Directorate (STMD) was allocated three payload opportunities on the EM-1 mission. These slots will be filled via the Centennial Challenges Program, NASA's flagship program for technology prize competitions, which directly engages the public, academia, and industry in open prize competitions to stimulate innovation. The NASA Science Mission Directorate (SMD) was allocated two payload opportunities on the EM-1 mission. The CubeSat Mission to Study Solar Particles (CuSP) payload will study the sources and acceleration mechanisms of solar and interplanetary particles in near-Earth orbit, support space weather research by determining proton radiation levels during Solar Energetic Particle (SEP) events and identifying suprathermal properties that could help predict geomagnetic storms. The LunaH-Map payload will help scientists understand the quantity of H-bearing materials in lunar cold traps (10 km), determine the concentration of H-bearing materials with 1m depth, and constrain the vertical distribution of H-bearing materials. The final three payload opportunities for the EM-1 mission were allocated for NASA's international space agency counterparts. The flight opportunities are intended to benefit the international space agency and NASA as well as further the collective space exploration goals. ArgoMoon is sponsored by ESA/ASI and will fly along with the ICPS on its disposal trajectory to perform proximity operations with the ICPS post-disposal, take external imagery of engineering and historical significance, and perform an optical communications demonstration. EQUULEUS, sponsored by JAXA, will fly to a libration orbit around the Earth-Moon L2 point and demonstrate trajectory control techniques within the Sun-Earth- Moon region for the first time by a nano spacecraft. The mission will also contribute to the future human exploration scenario by understanding the radiation environment in geospace and deep space, characterizing the flux of impacting meteors on the far side of the moon, and demonstrating the future deep space exploration scenario using the "deep space port" at Lagrange points. OMOTENASHI, also sponsored by JAXA, will land the smallest lunar lander to date on the lunar surface to demonstrate the feasibility of the hardware for distributed cooperative exploration system. Small landers will enable multi-point exploration, which is complimentary with large-scale human exploration. Once on the lunar surface, the OMOTENASHI spacecraft will observe the radiation and soil environments of the lunar surface by active radiation measurements and soil shear measurements. Following EM-1, Space Launch System will evolve to the more-powerful Block 1B configuration, which uses a new Exploration Upper Stage to increase the vehicle's LEO payload capability from 70 t to 105 t. With that transition, the Orion Stage Adapter, which will carry the secondary payloads on EM-1, will be phased out, and a new Universal Stage Adapter will be introduced, creating opportunities for flying larger secondary payloads. This paper will provide a brief status of SLS progress toward first launch; an overview of smallsat accommodations, integration, and operations on EM-1; information about the specific payloads flying on that launch; and a discussion of future accommodations and opportunities for secondary payloads on SLS for Exploration Mission-2 and beyond.

The SALSA Project - High-End Aerial 3d Camera

NASA Astrophysics Data System (ADS)

Rüther-Kindel, W.; Brauchle, J.

2013-08-01

The ATISS measurement drone, developed at the University of Applied Sciences Wildau, is an electrical powered motor glider with a maximum take-off weight of 25 kg including a payload capacity of 10 kg. Two 2.5 kW engines enable ultra short take-off procedures and the motor glider design results in a 1 h endurance. The concept of ATISS is based on the idea to strictly separate between aircraft and payload functions, which makes ATISS a very flexible research platform for miscellaneous payloads. ATISS is equipped with an autopilot for autonomous flight patterns but under permanent pilot control from the ground. On the basis of ATISS the project SALSA was undertaken. The aim was to integrate a system for digital terrain modelling. Instead of a laser scanner a new design concept was chosen based on two synchronized high resolution digital cameras, one in a fixed nadir orientation and the other in a oblique orientation. Thus from every object on the ground images from different view angles are taken. This new measurement camera system MACS-TumbleCam was developed at the German Aerospace Center DLR Berlin-Adlershof especially for the ATISS payload concept. Special advantage in comparison to laser scanning is the fact, that instead of a cloud of points a surface including texture is generated and a high-end inertial orientation system can be omitted. The first test flights show a ground resolution of 2 cm and height resolution of 3 cm, which underline the extraordinary capabilities of ATISS and the MACS measurement camera system.

SmallSat Spinning Lander with a Raman Spectrometer Payload for Future Ocean Worlds Exploration Missions

NASA Technical Reports Server (NTRS)

Ridenoure, R.; Angel, S. M.; Aslam, S.; Gorius, N.; Hewagama, T.; Nixon, C. A.; Sharma, S.

2017-01-01

We describe an Evolved Expendable Launch Vehicle Secondary Payload Adapter (ESPA)-class SmallSat spinning lander concept for the exploration of Europa or other Ocean World surfaces to ascertain the potential for life. The spinning lander will be ejected from an ESPA ring from an orbiting or flyby spacecraft and will carry on-board a standoff remote Spatial Heterodyne Raman spectrometer (SHRS) and a time resolved laser induced fluorescence spectrograph (TR-LIFS), and once landed and stationary the instruments will make surface chemical measurements. The SHRS and TR-LIFS have no moving parts have minimal mass and power requirements and will be able to characterize the surface and near-surface chemistry, including complex organic chemistry to constrain the ocean composition.

SmallSat Spinning Lander with a Raman Spectrometer Payload for Future Ocean Worlds Exploration Missions

NASA Astrophysics Data System (ADS)

Ridenoure, R.; Angel, S. M.; Aslam, S.; Gorius, N.; Hewagama, T.; Nixon, C. A.; Sharma, S.

2017-09-01

We describe an Evolved Expendable Launch Vehicle Secondary Payload Adapter (ESPA)-class SmallSat spinning lander concept for the exploration of Europa or other Ocean World surfaces to ascertain the potential for life. The spinning lander will be ejected from an ESPA ring from an orbiting or flyby spacecraft and will carry on-board a standoff remote Spatial Heterodyne Raman spectrometer (SHRS) and a time resolved laser induced fluorescence spectrograph (TR-LIFS), and once landed and stationary the instruments will make surface chemical measurements. The SHRS and TR-LIFS have no moving parts have minimal mass and power requirements and will be able to characterize the surface and near-surface chemistry, including complex organic chemistry to constrain the ocean composition.

Configuration development study of the X-24C hypersonic research airplane, phase 3

NASA Technical Reports Server (NTRS)

Combs, H. G.

1977-01-01

The conclusion evolved from the three phased study on the configuration development of the X-24C Hypersonic Research Airplane makes it evident that it is practical to design and build the high performance National Hypersonic Flight Research Facility airplane with today's state of the art within the cost and operational constraints established by NASA. The vehicle launched at 31.75 Mg from the B-52 can cruise for 40 seconds at Mach 6.78 on scramjets. Without scramjets it can approach Mach 8 with a 453.6 Kg payload or do 70 seconds of cruise at Mach 6 with a 2.27 Mg payload. Reduction in cost is possible with a vehicle scaled to a lesser mass and capability.

A reconfigurable robot with tensegrity structure using nylon artificial muscle

NASA Astrophysics Data System (ADS)

Wu, Lianjun; de Andrade, Monica Jung; Brahme, Tarang; Tadesse, Yonas; Baughman, Ray H.

2016-04-01

This paper describes the design and experimental investigation of a self-reconfigurable icosahedral robot for locomotion. The robot consists of novel and modular tensegrity structures, which can potentially maneuver in unstructured environments while carrying a payload. Twisted and Coiled Polymer (TCP) muscles were utilized to actuate the tensegrity structure as needed. The tensegrity system has rigid struts and flexible TCP muscles that allow keeping a payload in the central region. The TCP muscles provide large actuation stroke, high mechanical power per fiber mass and can undergo millions of highly reversible cycles. The muscles are electrothermally driven, and, upon stimulus, the heated muscles reconfigure the shape of the tensegrity structure. Here, we present preliminary experimental results that determine the rolling motion of the structure.

The large satellite program of ESA and its relevance for broadcast missions

NASA Astrophysics Data System (ADS)

Fromm, H.-H.; Herdan, B. L.

1981-03-01

In an investigation of the market prospects and payload requirements of future communications satellites, it was concluded that during the next 15 years many space missions will demand larger satellite platforms than those currently available. These platforms will be needed in connection with direct-broadcasting satellites, satellites required to enhance capacities in the case of traditional services, and satellites employed to introduce new types of satellite-based communications operating with small terminals. Most of the larger satellites would require the Ariane III capability, corresponding to about 1400 kg satellite mass in geostationary orbit. Attention is given to L-SAT platform capabilities and broadcast payload requirements, taking into account a European direct-broadcast satellite and Canadian direct-broadcast missions.

A pumped, two-phase flow heat transport system for orbiting instrument payloads

NASA Technical Reports Server (NTRS)

Fowle, A. A.

1981-01-01

A pumped two-phase (heat absorption/heat rejection) thermal transport system for orbiting instrument payloads is investigated. The thermofluid characteristics necessary for the system design are discussed. A preliminary design with a series arrangement of four instrument heat stations and six radiators in a single loop is described in detail, and the total mass is estimated to be 134 kg, with the radiators, instrument heat stations, and fluid reservoir accounting for approximately 86, 24, and 12 kg, respectively. The evaluation of preliminary test results shows that the system has potential advantages; however, further research is necessary in the areas of one-g and zero-g heat transfer coefficients/fluid regimes, fluid by-pass temperature control, and reliability of small pumps.

Japanese experiment module data management and communication system

NASA Astrophysics Data System (ADS)

Iizuka, Isao; Yamamoto, Harumitsu; Harada, Minoru; Eguchi, Iwao; Takahashi, Masami

The data management and communications system (DMCS) for the Japanese experiment module (JEM) being developed for the Space Station is described. Data generated by JEM experiments will be transmitted via TDRS (primary link) to the NASDA Operation Control Center. The DMSC will provide data processing, test and graphics handling, schedule planning support, and data display and facilitate subsystems, payloads, emergency operations, status, and diagnostics and healthchecks management. The ground segment includes a mainframe, mass storage, a workstation, and a LAN, with the capability of receiving and manipulating data from the JEM, the Space Station, and the payload. Audio and alert functions are also included. The DMCS will be connected to the interior of the module with through-bulkhead optical fibers.

Low gravity investigations in suborbital rockets

NASA Technical Reports Server (NTRS)

Wessling, Francis C.; Lundquist, Charles A.

1990-01-01

Two series of suborbital rocket missions are outlined which are intended to support materials and biotechnology investigations under microgravity conditions and enhance commercial rocket activity. The Consort series of missions employs the two-stage Starfire I rocket and recovery systems as well as a payload of three sealed or vented cylindrical sections. The Consort 1 and 2 missions are described which successfully supported six classes of experiments each. The Joust program is the second series of rocket missions, and the Prospector rocket is employed to provide comparable payload masses with twice as much microgravity time as the Consort series. The Joust and Consort missions provide 6-8 and 13-15 mins, respectively, of microgravity flight to support such experiments as polymer processing, scientific apparatus testing, and electrodeposition.

Deployment of the Telstar communications satellite

NASA Technical Reports Server (NTRS)

1984-01-01

Deployment of the Telstar communications satellite by the STS 41-D crew. The cylindrical spacecraft spins and rises past the shuttle stabilizer from its protective shield in the space shuttle payload bay. Behind it is another, closed protective cradle. Heavy clouds cover much of the water and land mass of earth in the background.

One vs two primary LOX feedline configuration study for the National Launch System

NASA Technical Reports Server (NTRS)

Dill, K.; Davis, D.; Bates, R.; Tarwater, R.

1992-01-01

Six single LOX feedline designs were evaluated for use on the National Launch Vehicle. A single feedline design, designated the 'Spider', was chosen and compared to the baseline system. The baseline configuration employs two 20-inch I.D. lines, each supplying LOX to three 650,000 lbf thrust Space Transportation Main Engines. Five single feedline diameters were examined for the spider configuration; 22, 24, 26, 28, and 30-inch I.D. System dry weights and LOX residuals were estimated. These parameters, along with calculated staged mass for the different single line and baseline configurations, were used to calculate the payload mass to orbit. For the cases where LOX is drained to minimum NPSP conditions, none of the single lines performed as well as the dual line system, although the 22-inch diameter single line compared well. However, for the cases where LOX is drained to operating levels (LOX level at the booster and spider manifolds for the dual and single line configurations, respectively), the 22 - 26-inch I.D. single line systems show a greater payload capability.

Conceptual design of an ascent-phase interceptor missile

DOE Office of Scientific and Technical Information (OSTI.GOV)

Salguero, D E

1994-11-01

A conceptual design for an air-launched interceptor missile to defend against theater ballistic missiles is presented. The missile is designed to intercept the target while ascending, during Or just after the boost phase, before it reaches exo-atmospheric flight. The interceptor consists of a two-stage booster and a shrouded kinetic-kill vehicle. This report concentrates on the booster design required to achieve reasonable standoff ranges. The kinetic-kill vehicle and shroud (the payload) is assumed to weigh 80 lb{sub m} (36 kg) and assumed to contain guidance computers for both the kill vehicle and the booster. The interceptor missile is about 6 mmore » long, .48 m in diameter and weighs about 900 kg. Allowing 25 sec for target detection, trajectory estimation, and interceptor launch, it can intercept 90 sec after target launch from a 220 km stand-off range at an altitude of 60 km. Trade-off studies show that the interceptor performance is most sensitive to the stage mass fractions (with the first-stage mass fraction the most important), the first-stage burn time and the payload weight.« less

The small satellite NINA-MITA to study galactic and solar cosmic rays in low-altitude polar orbit

NASA Astrophysics Data System (ADS)

Furano, G.; Bidoli, V.; Casolino, M.; de Pascale, M. P.; Iannucci, A.; Morselli, A.; Picozza, P.; Reali, E.; Sparvoli, R.; Bakaldin, A.; Galper, A.; Koldashov, M.; Korotkov, M.; Leonov, A.; Mikhailov, V.; Murashov, A.; Voronov, S.; Mazzenga, G.; Ricci, M.; Castellini, G.; Barbiellini, M.; Boezio, M.; Bonvicini, V.; Cirami, R.; Vacchi, A.; Zampa, N.; Ambriola, M.; Bellotti, R.; Cafagna, F.; Ciacio, F.; Circella, M.; de Marzo, C.; Adriani, O.; Papini, P.; Piccardi, S.; Spillantini, P.

The satellite MITA, carrying on board the scientific payload NINA-2, was launched on July the 15th, 2000 from the cosmodrome of Plesetsk (Russia) with a Cosmos-3M rocket. The satellite and the payload are currently operating within nominal parameters. NINA-2 is the first scientific payload for the technological flight of the Italian small satellite MITA. The detector used in this mission is identical to the one already flying on the Russian satellite Resurs-O1 n.4 in a 840-km sun-synchronous orbit, but makes use of the extensive computer and telemetry capabilities of MITA bus to improve the active data acquisition time. NINA physics objectives are to study cosmic nuclei from hydrogen to iron in the energy range between 10 MeV/n and 1 GeV/n during the years 2000-2003, that is the solar maximum period. The device is capable of charge identification up to iron with isotope sensitivity up to oxigen. The 87.3 degrees, 460 km altitude polar orbit allows investigations of cosmic rays of solar and galactic origin, so to study long and short term solar transient phenomena, and the study of the trapped radiation at higher geomagnetic cutoff.

Concept designs for NASA's Solar Electric Propulsion Technology Demonstration Mission

NASA Technical Reports Server (NTRS)

Mcguire, Melissa L.; Hack, Kurt J.; Manzella, David H.; Herman, Daniel A.

2014-01-01

Multiple Solar Electric Propulsion Technology Demonstration Mission were developed to assess vehicle performance and estimated mission cost. Concepts ranged from a 10,000 kilogram spacecraft capable of delivering 4000 kilogram of payload to one of the Earth Moon Lagrange points in support of future human-crewed outposts to a 180 kilogram spacecraft capable of performing an asteroid rendezvous mission after launched to a geostationary transfer orbit as a secondary payload. Low-cost and maximum Delta-V capability variants of a spacecraft concept based on utilizing a secondary payload adapter as the primary bus structure were developed as were concepts designed to be co-manifested with another spacecraft on a single launch vehicle. Each of the Solar Electric Propulsion Technology Demonstration Mission concepts developed included an estimated spacecraft cost. These data suggest estimated spacecraft costs of $200 million - $300 million if 30 kilowatt-class solar arrays and the corresponding electric propulsion system currently under development are used as the basis for sizing the mission concept regardless of launch vehicle costs. The most affordable mission concept developed based on subscale variants of the advanced solar arrays and electric propulsion technology currently under development by the NASA Space Technology Mission Directorate has an estimated cost of $50M and could provide a Delta-V capability comparable to much larger spacecraft concepts.

Effect of long-term exposure to Low Earth Orbit (LEO) space environment

NASA Technical Reports Server (NTRS)

Zimcik, D. G.

1987-01-01

Data obtained from components and materials from the Solar Maximum Mission satellite are presented and compared to data for similar materials obtained from the Advanced Composite Materials Exposure to Space Experiment (ACOMEX) flown on Shuttle mission STS-41G. In addition to evaluation of surface erosion and mass loss that may be of importance to very long-term missions, comparisons of solar absorptance and thermal emittance measurements for both long and short term exposures were made. Although the ratio of absorptance over emittance can be altered by proper choice of materials to ensure a proper operating environment for the spacecraft, once the thermal design is established, it is important that the material properties not change in order to maintain the operating environment for many payload and bus items such as electronics, batteries, fuel, etc. However, data presented show significant changes after short exposure in low Earth environment. Moreover, the measured changes are shown to differ according to the manner of exposure, i.e., normal or oblique, which also affects the resultant eroded surface morphology. These results identify constraints to be considered in development of flight experiments or laboratory testing.

EUROPA Multiple-Flyby Trajectory Design

NASA Technical Reports Server (NTRS)

Buffington, Brent; Campagnola, Stefano; Petropoulos, Anastassios

2012-01-01

As reinforced by the 2011 NRC Decadal Survey, Europa remains one of the most scientifically intriguing targets in planetary science due to its potential suitability for life. However, based on JEO cost estimates and current budgetary constraints, the Decadal Survey recommended-and later directed by NASA Headquarters-a more affordable pathway to Europa exploration be derived. In response, a flyby-only proof-of-concept trajectory has been developed to investigate Europa. The trajectory, enabled by employing a novel combination of new mission design techniques, successfully fulfills a set of Science Definition Team derived scientific objectives carried out by a notional payload including ice penetrating radar, topographic imaging, and short wavelength infrared observations, and ion neutral mass spectrometry in-situ measurements. The current baseline trajectory, referred to as 11-F5, consists of 34 Europa and 9 Ganymede flybys executed over the course of 2.4 years, reached a maximum inclination of 15 degrees, has a deterministic delta v of 157 m/s (post-PJR), and has a total ionizing dose of 2.06 Mrad (Si behind 100 mil Al, spherical shell). The 11-F5 trajectory and more generally speaking, flyby-only trajectories-exhibit a number of potential advantages over an Europa orbiter mission.

Evidence for a maximum mass cut-off in the neutron star mass distribution and constraints on the equation of state

NASA Astrophysics Data System (ADS)

Alsing, Justin; Silva, Hector O.; Berti, Emanuele

2018-07-01

We infer the mass distribution of neutron stars in binary systems using a flexible Gaussian mixture model and use Bayesian model selection to explore evidence for multimodality and a sharp cut-off in the mass distribution. We find overwhelming evidence for a bimodal distribution, in agreement with previous literature, and report for the first time positive evidence for a sharp cut-off at a maximum neutron star mass. We measure the maximum mass to be 2.0 M⊙ < mmax < 2.2 M⊙ (68 per cent), 2.0 M⊙ < mmax < 2.6 M⊙ (90 per cent), and evidence for a cut-off is robust against the choice of model for the mass distribution and to removing the most extreme (highest mass) neutron stars from the data set. If this sharp cut-off is interpreted as the maximum stable neutron star mass allowed by the equation of state of dense matter, our measurement puts constraints on the equation of state. For a set of realistic equations of state that support >2 M⊙ neutron stars, our inference of mmax is able to distinguish between models at odds ratios of up to 12:1, whilst under a flexible piecewise polytropic equation-of-state model our maximum mass measurement improves constraints on the pressure at 3-7× the nuclear saturation density by ˜ 30-50 per cent compared to simply requiring mmax > 2 M⊙. We obtain a lower bound on the maximum sound speed attained inside the neutron star of c_ s^max > 0.63c (99.8 per cent), ruling out c_ s^max < c/√{3} at high significance. Our constraints on the maximum neutron star mass strengthen the case for neutron star-neutron star mergers as the primary source of short gamma-ray bursts.

Acoustic Test Results of Melamine Foam with Application to Payload Fairing Acoustic Attenuation Systems

NASA Technical Reports Server (NTRS)

Hughes, William O.; McNelis, Anne M.

2014-01-01

A spacecraft at launch is subjected to a harsh acoustic and vibration environment resulting from the passage of acoustic energy, created during the liftoff of a launch vehicle, through the vehicle's payload fairing. In order to ensure the mission success of the spacecraft it is often necessary to reduce the resulting internal acoustic sound pressure levels through the usage of acoustic attenuation systems. Melamine foam, lining the interior walls of the payload fairing, is often utilized as the main component of such a system. In order to better understand the acoustic properties of melamine foam, with the goal of developing improved acoustic attenuation systems, NASA has recently performed panel level testing on numerous configurations of melamine foam acoustic treatments at the Riverbank Acoustical Laboratory. Parameters assessed included the foam's thickness and density, as well as the effects of a top outer cover sheet material and mass barriers embedded within the foam. This testing followed the ASTM C423 standard for absorption and the ASTM E90 standard for transmission loss. The acoustic test data obtained and subsequent conclusions are the subjects of this paper.

Advanced telemetry systems for payloads. Technology needs, objectives and issues

NASA Technical Reports Server (NTRS)

1990-01-01

The current trends in advanced payload telemetry are the new developments in advanced modulation/coding, the applications of intelligent techniques, data distribution processing, and advanced signal processing methodologies. Concerted efforts will be required to design ultra-reliable man-rated software to cope with these applications. The intelligence embedded and distributed throughout various segments of the telemetry system will need to be overridden by an operator in case of life-threatening situations, making it a real-time integration issue. Suitable MIL standards on physical interfaces and protocols will be adopted to suit the payload telemetry system. New technologies and techniques will be developed for fast retrieval of mass data. Currently, these technology issues are being addressed to provide more efficient, reliable, and reconfigurable systems. There is a need, however, to change the operation culture. The current role of NASA as a leader in developing all the new innovative hardware should be altered to save both time and money. We should use all the available hardware/software developed by the industry and use the existing standards rather than inventing our own.

Benefits of Nuclear Electric Propulsion for Outer Planet Exploration

NASA Technical Reports Server (NTRS)

Kos, Larry; Johnson, Les; Jones, Jonathan; Trausch, Ann; Eberle, Bill; Woodcock, Gordon; Brady, Hugh J. (Technical Monitor)

2002-01-01

Nuclear electric propulsion (NEP) offers significant benefits to missions for outer planet exploration. Reaching outer planet destinations, especially beyond Jupiter, is a struggle against time and distance. For relatively near missions, such as a Europa lander, conventional chemical propulsion and NEP offer similar performance and capabilities. For challenging missions such as a Pluto orbiter, neither chemical nor solar electric propulsion are capable while NEP offers acceptable performance. Three missions are compared in this paper: Europa lander, Pluto orbiter, and Titan sample return, illustrating how performance of conventional and advanced propulsion systems vary with increasing difficulty. The paper presents parametric trajectory performance data for NEP. Preliminary mass/performance estimates are provided for a Europa lander and a Titan sample return system, to derive net payloads for NEP. The NEP system delivers payloads and ascent/descent spacecraft to orbit around the target body, and for sample return, delivers the sample carrier system from Titan orbit to an Earth transfer trajectory. A representative scientific payload 500 kg was assumed, typical for a robotic mission. The resulting NEP systems are 100-kWe class, with specific impulse from 6000 to 9000 seconds.

A Compact Tandem Two-Step Laser Time-of-Flight Mass Spectrometer for In Situ Analysis of Non-Volatile Organics on Planetary Surfaces

NASA Technical Reports Server (NTRS)

Getty, Stephanie A.; Brinckerhoff, William B.; Li, Xiang; Elsila, Jamie; Cornish, Timothy; Ecelberger, Scott; Wu, Qinghao; Zare, Richard

2014-01-01

Two-step laser desorption mass spectrometry is a well suited technique to the analysis of high priority classes of organics, such as polycyclic aromatic hydrocarbons, present in complex samples. The use of decoupled desorption and ionization laser pulses allows for sensitive and selective detection of structurally intact organic species. We have recently demonstrated the implementation of this advancement in laser mass spectrometry in a compact, flight-compatible instrument that could feasibly be the centerpiece of an analytical science payload as part of a future spaceflight mission to a small body or icy moon.

Hermod: optical payload technology demonstrator flying on PROBA-V: overview of the payload development, testing and results after 1 year in orbit exploitation

NASA Astrophysics Data System (ADS)

Hernandez, S.; Blasco, J.; Henriksen, V.; Samuelsson, H.; Navasquillo, O.; Grimsgaard, M.; Mellab, K.

2017-11-01

Proba-V is the third mission of ESA's Programme for In-orbit Technology Demonstration (IOD), based on a small, high performance satellite platform and a compact payload. Besides, the main satellite instrument aiming at Vegetation imaging, Proba-V embarks five technological payloads providing early flight opportunities for novel instruments and space technologies. Successfully launched by the ESA VEGA launcher in May 2013, it has now completed its commissioning and the full calibration of platform, main instrument and additional payloads and is, since last October, fully operational. The High dEnsity space foRM cOnnector Demonstration or HERMOD is the last payload selected to fly on Proba-V. The payload objective is to validate through an actual launch and in orbit high-density optical fibre cable assembly, cumulate space heritage for fibre optics transmission and evaluate possible degradation induced by the space environment compared to on-ground tests. The future applications of this technology are for intrasatellite optical communications in view of mass reduction, the electrical grounding simplification and to increase the transmission rate. The project has been supported under an ESA GSTP contract. T&G Elektro (Norway) developed and tested the different optical cable assembly to be validated in the payload. The electrooptic modules, control, power and mechanical interfaces have been developed by DAS Photonics (Spain). The payload contains four optical channels to be studied through the experiment, two assemblies with MTP/PC connectors and two assemblies with MPO/APC connectors. Optical data is transmitted in the four independent channels using two optoelectronic conversion modules (SIOS) working at 100Mbps including 2 full duplex channels each. A FPGA is used to generate, receive and compare the different binary patterns. The number of errors (if any) and Bit Error Rate (BER) is sent to the satellite TM interface. HERMOD successfully went through all mechanical and environmental tests before the integration in a very limited time. The telemetry data is currently sent to ground on daily basis. All the channels have survived the launch and no BER has been measured with the exception of channel 2, currently recording a BER of 3.06*10-16, that exhibits from time to time a burst of errors due to synchronizing issues of the initial data frame. It is expected to observe during the operating life of the payload the first errors within the channel 4 which was designed on purpose with reduced power margin. This paper will present the full overview of the HERMOD technology demonstrator including the development, testing, validation activity, integration, commissioning and 1 year in-orbit exploitation results.

Power Systems Evaluated for Solar Electric Propulsion Vehicles

NASA Technical Reports Server (NTRS)

Kerslake, Thomas W.; Gefert, Leon P.

2000-01-01

Solar electric propulsion (SEP) mission architectures are applicable to a wide range NASA missions including the robotic exploration of the outer planets in the next decade and the human exploration of Mars within the next 2 decades. SEP enables architectures that are very mass efficient with reasonable power levels (1-MW class) aerobrake and cryogenic upper-stage transportation technologies are utilized. In this architecture, the efficient SEP stage transfers the payload from low Earth orbit (LEO) High Energy Elliptical Parking Orbit (HEEPO) within a period of 6 to 12 months. highthrust, cryogenic upper stage and payload then separate from the SEP vehicle for injection to the planetary target, allowing for fast heliocentric trip times. This mission architecture offers a potential reduction in mass to LEO in comparison to alternative all-chemical nuclear propulsion schemes. Mass reductions may allow launch vehicle downsizing enable missions that would have been grounded because of cost constraints. The preceding figure illustrates a conceptual SEP stage design for a human Mars mission. Researchers at the NASA Glenn Research Center at Lewis Field designed conceptual SEP vehicle, conceived the mission architecture to use this vehicle, and analyzed the vehicle s performance. This SEP stage has a dry mass of 35 metric tons (MT), 40 MT of xenon propellant, and a photovoltaic array that spans 110 m, providing power to a cluster of eight 100-kW Hall thrusters. The stage can transfer an 80-MT payload and upper stage to the desired HEEPO. Preliminary packaging studies show this space-station-class SEP vehicle meets the proposed "Magnum" launch vehicle and volume requirements with considerable margin. An SEP vehicle for outer planetary missions, such as the Europa Mapper Mission, would be dramatically smaller than human Mars mission SEP stage. In this mission architecture, the SEP power system with the payload to provide spacecraft power throughout the mission. Several photovoltaic array design concepts were considered for the SEP vehicle power system for the human mission to Mars. These include a space station derivative, a SCARLET (Solar Concentrator Arrays with Refractive Linear Element Technology) derivative, and a hybrid inflatable-deployable thin polymer membrane array with thin-film solar cells (as shown in the concept illustration). This concept is based on a design developed for the Next Generation Space Telescope Sun shield. The array is divided into 16 independent electrical sections with 500-V, negative-grounded solar cell strings. The power system employs a channelized, 500-Vdc power management and distribution (PMAD) architecture with lithium ion batteries for energy storage for vehicle and payload secondary loads (the high-power Hall thrusters do not operate in eclipse periods). The 500-V PMAD voltage permits "direct-drive" thruster operation, greatly reducing the power processing unit size, complexity, and power loss. Similar power system architecture, designs, and technology are assumed for the Europa Mapper Mission SEP vehicle. The primary exceptions are that the photovoltaic array is assumed to consist of two rectangular wings and that the power system rating is 15 kW in Earth orbit and 200 W at Europa. To size the SEP vehicle power system, a dedicated Fortran code was developed to predict detailed power system performance, mass, and thermal control requirements. This code also modeled all the relevant Earth orbit environments; that is, the particulate radiation, plasma, meteoroids and debris, ultraviolet radiation, contamination, and thermal conditions. Analysis results for the Human Mars Mission SEP vehicle show a power system mass of 9-MT and photovoltaic array area of 5800-square meters for the thin-membrane design concept with CuInS2 thin-film cells. Power processing unit input power for a thin-membrane array design with three-junction, amorphous SiGe solar cells is shown in the graph. Power falls off rapidly inhe first weeks of the mission because of light-induced (Staebler-Wronksi) solar cell losses. During the next 200 days, power decreases steadily as the SEP stage spirals through the proton belts and sustains the bulk of the mission radiation damage. Once the vehicle apogee is above approximately four Earth radii, little additional degradation is incurred. From 400 to 800 days, a 1100-km "parking" orbit is maintained to await the next payload transfer opportunity. This orbit is below the main proton belt, and thus, little radiation dose is accumulated during this time period. During the second LEO-to-HEEPO transfer, power degrades somewhat further, but power requirements are still met. In comparison, the Europa Mapper SEP vehicle power system had a mass of 150 kg and a thin membrane array area of 100 square meters.

Retrievable payload carrier, next generation Long Duration Exposure Facility: Update 1992

NASA Technical Reports Server (NTRS)

Perry, A. T.; Cagle, J. A.; Newman, S. C.

1993-01-01

Access to space and cost have been two major inhibitors of low Earth orbit research. The Retrievable Payload Carrier (RPC) Program is a commercial space program which strives to overcome these two barriers to space experimentation. The RPC Program's fleet of spacecraft, ground communications station, payload processing facility, and experienced integration and operations team will provide a convenient 'one-stop shop' for investigators seeking to use the unique vantage point and environment of low Earth orbit for research. The RPC is a regularly launched and retrieved, free-flying spacecraft providing resources adequate to meet modest payload/experiment requirements, and presenting ample surface area, volume, mass, and growth capacity for investigator usage. Enhanced capabilities of ground communications, solar-array-supplied electrical power, central computing, and on-board data storage pick up on the path where NASA's Long Duration Exposure Facility (LDEF) blazed the original technology trail. Mission lengths of 6-18 months, or longer, are envisioned. The year 1992 was designated as the 'International Space Year' and coincides with the 500th anniversary of Christopher Columbus's voyage to the New World. This is a fitting year in which to launch the full scale development of our unique shop of discovery whose intent is to facilitate retrieving technological rewards from another new world: space. Presented is an update on progress made on the RPC Program's development since the November 1991 LDEF Materials Workshop.

A modular design for rapid-response telecoms and navigation missions

NASA Astrophysics Data System (ADS)

Davies, P.; Liddle, D.; Buckley, John; Sweeting, M.; Roussel-Dupre, Diane; Caffrey, Michael

2004-11-01

Surrey Satellite Technology Ltd and Los Alamos National Laboratory are together building the Cibola Flight Experiment (CFESat), a mission with the aim of flight-proving a reconfigurable processor payload intended for a Low Earth Orbit system. The mission will survey portions of the VHF and UHF radio spectra. The satellite will be launched by the Space Test Program in September 2006 on the USAF Evolved Expendable Launch Vehicle (EELV) using the EELV's Secondary Payload Adapter (ESPA) that allows up to six small satellites to be launched as "piggyback" passengers with larger spacecraft. The payload is based on networks of reprogrammable, Field Programmable Gate Arrays (FPGAs) to process the received signals for ionospheric and lightning studies. The objective is to validate the on-orbit use of commercial, reconfigurable FPGA technology utilizing several different single-event upset mitigation schemes. It will also detect and measure impulsive events that occur in a complex background. SSTL's satellite platform is based on a new, ESPA- compatible, structure housing subsystems and equipments with proven flight heritage from SSTL's disaster monitoring constellation (DMC) and the Topsat mission satellite due for launch in 2005. The structure is mechanically quite complex for a microsatellite having both deployed solar panels and a pair of long booms as part of the payload. The satellite design is highly constrained by the mass and volume requirements of the EELV/EPSA.

GEP, A Geophysical and Environemental integrated payload for ExoMars

NASA Astrophysics Data System (ADS)

Spohn, T.; Lognonne, P.; Dehant, V.; Giardini, D.; Friis-Christensen, E.; Calcutt, S.; GEP Team

The goal of the GEP proposed onboard the ExoMars mission is to provide the first complete set of geophysical and environmental data of Mars. A full mass of 20 kg is envisaged, enabling a payload of about 5 kg serviced by common integrated subsystems. GEP will first monitor the present Martian climate and meteorology by providing a unique monitoring on potential hazards for future human exploration missions (radiations, atmospheric electricity, dust) and on atmospheric parameters (wind, pressure, temperature, humidity). Such a long term monitoring has never been performed since the Viking landers. GEP will then provide, for the first time, a complete geophysical monitoring of Mars. It will search for remote and regional seismic activity, will measure the heat flux of the planets, will monitor the rotation of Mars and will study the magnetic field at the surface and finally will constrain the subsurface in the vicinity of the ExoMars landing site and the deep interior. By providing these new geophysical data and associated constraints on the interior and on the actual geologic activity of the surface, GEP will provide a major step in our understanding of the geological evolution of the planet and the habitability conditions during the first billion years, enabling a full understanding of the surface and mineralogical observations performed by the Pasteur payload onboard the ExoMars rover and by the payload onboard the MSL NASA 2009 mission.

NH11B-1726: FrankenRaven: A New Platform for Remote Sensing

NASA Technical Reports Server (NTRS)

Dahlgren, Robert; Fladeland, Matthew M.; Pinsker, Ethan A.; Jasionowicz, John P.; Jones, Lowell L.; Pscheid, Matthew J.

2016-01-01

Small, modular aircraft are an emerging technology with a goal to maximize flexibility and enable multi-mission support. This reports the progress of an unmanned aerial system (UAS) project conducted at the NASA Ames Research Center (ARC) in 2016. This interdisciplinary effort builds upon the success of the 2014 FrankenEye project to apply rapid prototyping techniques to UAS, to develop a variety of platforms to host remote sensing instruments. In 2016, ARC received AeroVironment RQ-11A and RQ-11B Raven UAS from the US Department of the Interior, Office of Aviation Services. These aircraft have electric propulsion, a wingspan of roughly 1.3m, and have demonstrated reliability in challenging environments. The Raven airframe is an ideal foundation to construct more complex aircraft, and student interns using 3D printing were able to graft multiple Raven wings and fuselages into FrankenRaven aircraft. Aeronautical analysis shows that the new configuration has enhanced flight time, payload capacity, and distance compared to the original Raven. The FrankenRaven avionics architecture replaces the mil-spec avionics with COTS technology based upon the 3DR Pixhawk PX4 autopilot with a safety multiplexer for failsafe handoff to 2.4 GHz RC control and 915 MHz telemetry. This project demonstrates how design reuse, rapid prototyping, and modular subcomponents can be leveraged into flexible airborne platforms that can host a variety of remote sensing payloads and even multiple payloads. Modularity advances a new paradigm: mass-customization of aircraft around given payload(s). Multi-fuselage designs are currently under development to host a wide variety of payloads including a zenith-pointing spectrometer, a magnetometer, a multi-spectral camera, and a RGB camera. After airworthiness certification, flight readiness review, and test flights are performed at Crows Landing airfield in central California, field data will be taken at Kilauea volcano in Hawaii and other locations.

FrankenRaven: A New Platform for Remote Sensing

NASA Astrophysics Data System (ADS)

Dahlgren, R. P.; Fladeland, M. M.; Pinsker, E. A.; Jasionowicz, J. P.; Jones, L. L.; Mosser, C. D.; Pscheid, M. J.; Weidow, N. L.; Kelly, P. J.; Kern, C.; Werner, C. A.; Johnson, M. S.

2016-12-01

Small, modular aircraft are an emerging technology with a goal to maximize flexibility and enable multi-mission support. This reports the progress of an unmanned aerial system (UAS) project conducted at the NASA Ames Research Center (ARC) in 2016. This interdisciplinary effort builds upon the success of the 2014 FrankenEye project to apply rapid prototyping techniques to UAS, to develop a variety of platforms to host remote sensing instruments. In 2016, ARC received AeroVironment RQ-11A and RQ-11B Raven UAS from the US Department of the Interior, Office of Aviation Services. These aircraft have electric propulsion, a wingspan of roughly 1.3m, and have demonstrated reliability in challenging environments. The Raven airframe is an ideal foundation to construct more complex aircraft, and student interns using 3D printing were able to graft multiple Raven wings and fuselages into "FrankenRaven" aircraft. Aeronautical analysis shows that the new configuration has enhanced flight time, payload capacity, and distance compared to the original Raven. The FrankenRaven avionics architecture replaces the mil-spec avionics with COTS technology based upon the 3DR Pixhawk PX4 autopilot with a safety multiplexer for failsafe handoff to 2.4 GHz RC control and 915 MHz telemetry. This project demonstrates how design reuse, rapid prototyping, and modular subcomponents can be leveraged into flexible airborne platforms that can host a variety of remote sensing payloads and even multiple payloads. Modularity advances a new paradigm: mass-customization of aircraft around given payload(s). Multi-fuselage designs are currently under development to host a wide variety of payloads including a zenith-pointing spectrometer, a magnetometer, a multi-spectral camera, and a RGB camera. After airworthiness certification, flight readiness review, and test flights are performed at Crows Landing airfield in central California, field data will be taken at Kilauea volcano in Hawaii and other locations.

Aesops Guide to Space Acquisition: Glimpses of Past and Future

DTIC Science & Technology

2015-04-01

Defense AT&L: March–April 2015 6 Aesop’s Guide to Space Acquisition Glimpses of Past and Future John Krieger n William “Lance” Krieger n Rick...U.S.C.) § 2273—Policy regarding assured access to space : national security payloads, establishes the requirement for the President to ensure our...access to space . (a) Policy—It is the policy of the United States for the President to undertake actions appropriate to ensure, to the maximum

40 CFR 75.72 - Determination of NOX mass emissions for common stack and multiple stack configurations.

Code of Federal Regulations, 2012 CFR

2012-07-01

... the affected units as the difference between NOX mass emissions measured in the common stack and NOX... emissions using the maximum potential NOX emission rate, the maximum potential flow rate, and either the maximum potential CO2 concentration or the minimum potential O2 concentration (as applicable). The maximum...

Trajectory optimization for A S.S.T.O. using in-flight LOX collection

NASA Astrophysics Data System (ADS)

Saint-Mard, M.; Hendrick, P.

A key point for a space mission (launch of a satellite, earth observation,…) is the optimization of the vehicle trajectory in order to burn the smallest quantity of propelant and then maximize the payload. This is true for evay space vehicle, but especially it is a crucial point for a Single-Stage-To-Orbit (SSTO) where the choice of a bad trajectory can result in an unrealizable vehicle due to the large airbreathing part of the flight In this study, we discuss the trajectory optimization for a Vertical Take-Off and Horizontal Landing (VTOHL) SSTO using supersonic in-flight atmospheric oxygen collection during a cruise phase (constant speed & constant altitude). This collected oxygen is stored in the LOX tanks and reused in the final rocket phase. This SSTO bas a Blended Body aerodynamic configuration as the one chosen by Lockheed Martin for its new space launcher (VentureStar and X-33). This SSTO uses rocket engines from take-off to Mach 1.7 and also for the exoatmospheric flight phase (that means for an altitude higher than 30km and a Mach number evolution from 6.8 to about 20). Between these two rocket phases, the SSTO is propelled by a subsonic ramjet. To perform this study, we use 2 computer programs (running on a home Computer): the first one allows to estimate the SSTO performances (TOGW, dry weight, hydrogen and oxygen consumptions) for a fixed payload mass and the second one permits the evaluation of the payload mass for a fixed TOGW.

Using NASA's Space Launch System to Enable Game Changing Science Mission Designs

NASA Technical Reports Server (NTRS)

Creech, Stephen D.

2013-01-01

NASA's Marshall Space Flight Center is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will help restore U.S. leadership in space by carrying the Orion Multi-Purpose Crew Vehicle and other important payloads far beyond Earth orbit. Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids, Mars, and the outer solar system. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required with several gravity-assist planetary fly-bys to achieve the necessary outbound velocity. The SLS rocket, using significantly higher C3 energies, can more quickly and effectively take the mission directly to its destination, reducing trip times and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as monolithic telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

Project Apophis for integrated research of minor body of the Solar System

NASA Astrophysics Data System (ADS)

Shustov, Boris M.; Martynov, Maxim; Zakharov, Alexander; Simonov, Alexander; Pol, Vadim

The results of pre-Phase A study of the project of space mission to a minor body of the Solar System are described. For definiteness of design the famous asteroid Apophis was chosen. This hectometer size asteroid is selected as a typical potentially hazardous minor body. The study was performed in cooperation of institutes of the Russian Academy of Sciences (Space Research Institute and Institute of Astronomy) and Roscosmos (Lavochkin Association). The major goals of the mission preliminary named “Apophis” are to carry out a study of physical and chemical properties of a potentially hazardous asteroid and to put a special radio beacon into circum-asteroid orbit aimed to precise determination of the asteroid’s orbital parameters. The time schedule is connected to the upcoming close encounter of Apophis with the Earth in 2029. The period around 2020 seems to be the most reasonable window for the launch. Selection of the launch date should meet the following optimization requirements: • minimal summary velocity consumption; • maximal mass of the SC on orbit of Apophis; • favorable conditions for observation of asteroid from the Earth. General features of a mission to Apophis are described. The total mass of payload is about 800 kg. Both distant and contact (if a lander option will be included) mode of study are planned. The expected lifetime of the mission is about 5 years (10 years for the beacon). The unique science instrument for the Apophis mission is a beacon itself and system of precise registration of position and velocity of the spacecraft that uses the beacon. Besides practical things these will provide fine data for study dynamical effects of the motion of minor body in the Solar System. The science instruments designed for study of bulk characteristics of the asteroid, its internal structure, properties of regolith and exosphere are included in the payload. Most of them are analogues of the instruments included in the “Phobos-Grunt” mission. The maximum use of the heritage (findings) of the mission “Phobos-Grunt” is an important feature that reduces the cost of the project Finally, we consider possible international cooperation on ground segment. It represents the capacity utilization of the VLBI (in particular opportunities of JIVE) for the trajectory measurements.

Design of a fast Mars space transfer system

NASA Astrophysics Data System (ADS)

Woo, Henry H.; Glass, James F.; Roy, Claude

1992-02-01

Architecture strategies and concepts for manned missions to Mars are being developed by NASA and industry. This paper addresses the key Mars transfer vehicle (MTV) design requirements which include surface payload mass, MTV mass, propulsion system characteristics, launch vehicle capability, in-space operations, abort considerations, crew exposure to interplanetary environments, and crew reconditioning for planetary entry. Different mission strategies are presented along with their implications. A representative artificial-g MTV using nuclear thermal propulsion is defined to show concepts which minimize extravehicular activity operations for in-space assembly, inspection, and maintenance.

Optimal trajectories for hypersonic launch vehicles

NASA Technical Reports Server (NTRS)

Ardema, Mark D.; Bowles, Jeffrey V.; Whittaker, Thomas

1992-01-01

In this paper, we derive a near-optimal guidance law for the ascent trajectory from Earth surface to Earth orbit of a hypersonic, dual-mode propulsion, lifting vehicle. Of interest are both the optimal flight path and the optimal operation of the propulsion system. The guidance law is developed from the energy-state approximation of the equations of motion. The performance objective is a weighted sum of fuel mass and volume, with the weighting factor selected to give minimum gross take-off weight for a specific payload mass and volume.

MACSAT - A Near Equatorial Earth Observation Mission

NASA Astrophysics Data System (ADS)

Kim, B. J.; Park, S.; Kim, E.-E.; Park, W.; Chang, H.; Seon, J.

MACSAT mission was initiated by Malaysia to launch a high-resolution remote sensing satellite into Near Equatorial Orbit (NEO). Due to its geographical location, Malaysia can have large benefits from NEO satellite operation. From the baseline circular orbit of 685 km altitude with 7 degrees of inclination, the neighboring regions around Malaysian territory can be frequently monitored. The equatorial environment around the globe can also be regularly observed with unique revisit characteristics. The primary mission objective of MACSAT program is to develop and validate technologies for a near equatorial orbit remote sensing satellite system. MACSAT is optimally designed to accommodate an electro-optic Earth observation payload, Medium-sized Aperture Camera (MAC). Malaysian and Korean joint engineering teams are formed for the effective implementation of the satellite system. An integrated team approach is adopted for the joint development for MACSAT. MAC is a pushbroom type camera with 2.5 m of Ground Sampling Distance (GSD) in panchromatic band and 5 m of GSD in four multi-spectral bands. The satellite platform is a mini-class satellite. Including MAC payload, the satellite weighs under 200 kg. Spacecraft bus is designed optimally to support payload operations during 3 years of mission life. The payload has 20 km of swath width with +/- 30 o of tilting capability. 32 Gbits of solid state recorder is implemented as the mass image storage. The ground element is an integrated ground station for mission control and payload operation. It is equipped with S- band up/down link for commanding and telemetry reception as well as 30 Mbps class X-band down link for image reception and processing. The MACSAT system is capable of generating 1:25,000-scale image maps. It is also anticipated to have capability for cross-track stereo imaging for Digital elevation Model (DEM) generation.

Biomechanical Assessment of Rucksack Shoulder Strap Attachment Location: Effect on Load Distribution to the Torso

DTIC Science & Technology

2001-05-01

varied vertical heights and horizontal positions. Optotrak position markers indicated by white diamonds). The position and mass of the payload (25 kg...shoulders and 45 N (+/- 2) at the waist and recorded for all configurations. An Optotrak 3-D motion tracking system, accuracy +/- 0.1 mm, was used to

The ultimate limits of the relativistic rocket equation. The Planck photon rocket

NASA Astrophysics Data System (ADS)

Haug, Espen Gaarder

2017-07-01

In this paper we look at the ultimate limits of a photon propulsion rocket. The maximum velocity for a photon propulsion rocket is just below the speed of light and is a function of the reduced Compton wavelength of the heaviest subatomic particles in the rocket. We are basically combining the relativistic rocket equation with Haug's new insight on the maximum velocity for anything with rest mass. An interesting new finding is that in order to accelerate any subatomic "fundamental" particle to its maximum velocity, the particle rocket basically needs two Planck masses of initial load. This might sound illogical until one understands that subatomic particles with different masses have different maximum velocities. This can be generalized to large rockets and gives us the maximum theoretical velocity of a fully-efficient and ideal rocket. Further, no additional fuel is needed to accelerate a Planck mass particle to its maximum velocity; this also might sound absurd, but it has a very simple and logical solution that is explained in this paper.

Human Mars Entry, Descent, and Landing Architecture Study Overview

NASA Technical Reports Server (NTRS)

Cianciolo, Alicia D.; Polsgrove, Tara T.

2016-01-01

The Entry, Descent, and Landing (EDL) Architecture Study is a multi-NASA center activity to analyze candidate EDL systems as they apply to human Mars landing in the context of the Evolvable Mars Campaign. The study, led by the Space Technology Mission Directorate (STMD), is performed in conjunction with the NASA's Science Mission Directorate and the Human Architecture Team, sponsored by NASA's Human Exploration and Operations Mission Directorate. The primary objective is to prioritize future STMD EDL technology investments by (1) generating Phase A-level designs for selected concepts to deliver 20 t human class payloads, (2) developing a parameterized mass model for each concept capable of examining payloads between 5 and 40 t, and (3) evaluating integrated system performance using trajectory simulations. This paper summarizes the initial study results.

Orbiter entry trajectory corridors: 32000 pound payload, 67.5 percent center of gravity. [glide path data compilation

NASA Technical Reports Server (NTRS)

Treybig, J. H.

1975-01-01

Thermal and equilibrium glide boundaries were used to analyze and/or design shuttle orbiter entry trajectories. Plots are presented of orbiter thermal and equilibrium glide boundaries in the drag/mass-relative velocity dynamic pressure-relative velocity, and altitude-relative velocity planes for an orbiter having a 32,000 pound payload and a 67.5% center of gravity location. These boundaries were defined for control points 1 through 4 of the shuttle orbiter for 40 deg-30 deg and 38 deg-28 deg ramped angle of attack entry profiles and 40 deg, 38 deg, 35 deg, 30 deg, 28 deg, and 25 deg constant angle of attack entry profiles each at 20 deg, 15 deg, and 10 deg constant body flap settings.

Earth Science

NASA Image and Video Library

1992-07-18

Workers at Launch Complex 17 Pad A, Kennedy Space Center (KSC) encapsulate the Geomagnetic Tail (GEOTAIL) spacecraft (upper) and attached payload Assist Module-D upper stage (lower) in the protective payload fairing. GEOTAIL project was designed to study the effects of Earth's magnetic field. The solar wind draws the Earth's magnetic field into a long tail on the night side of the Earth and stores energy in the stretched field lines of the magnetotail. During active periods, the tail couples with the near-Earth magnetosphere, sometimes releasing energy stored in the tail and activating auroras in the polar ionosphere. GEOTAIL measures the flow of energy and its transformation in the magnetotail and will help clarify the mechanisms that control the imput, transport, storage, release, and conversion of mass, momentum, and energy in the magnetotail.

A Mars Sample Return Sample Handling System

NASA Technical Reports Server (NTRS)

Wilson, David; Stroker, Carol

2013-01-01

We present a sample handling system, a subsystem of the proposed Dragon landed Mars Sample Return (MSR) mission [1], that can return to Earth orbit a significant mass of frozen Mars samples potentially consisting of: rock cores, subsurface drilled rock and ice cuttings, pebble sized rocks, and soil scoops. The sample collection, storage, retrieval and packaging assumptions and concepts in this study are applicable for the NASA's MPPG MSR mission architecture options [2]. Our study assumes a predecessor rover mission collects samples for return to Earth to address questions on: past life, climate change, water history, age dating, understanding Mars interior evolution [3], and, human safety and in-situ resource utilization. Hence the rover will have "integrated priorities for rock sampling" [3] that cover collection of subaqueous or hydrothermal sediments, low-temperature fluidaltered rocks, unaltered igneous rocks, regolith and atmosphere samples. Samples could include: drilled rock cores, alluvial and fluvial deposits, subsurface ice and soils, clays, sulfates, salts including perchlorates, aeolian deposits, and concretions. Thus samples will have a broad range of bulk densities, and require for Earth based analysis where practical: in-situ characterization, management of degradation such as perchlorate deliquescence and volatile release, and contamination management. We propose to adopt a sample container with a set of cups each with a sample from a specific location. We considered two sample cups sizes: (1) a small cup sized for samples matching those submitted to in-situ characterization instruments, and, (2) a larger cup for 100 mm rock cores [4] and pebble sized rocks, thus providing diverse samples and optimizing the MSR sample mass payload fraction for a given payload volume. We minimize sample degradation by keeping them frozen in the MSR payload sample canister using Peltier chip cooling. The cups are sealed by interference fitted heat activated memory alloy caps [5] if the heating does not affect the sample, or by crimping caps similar to bottle capping. We prefer cap sealing surfaces be external to the cup rim to prevent sample dust inside the cups interfering with sealing, or, contamination of the sample by Teflon seal elements (if adopted). Finally the sample collection rover, or a Fetch rover, selects cups with best choice samples and loads them into a sample tray, before delivering it to the Earth Return Vehicle (ERV) in the MSR Dragon capsule as described in [1] (Fig 1). This ensures best use of the MSR payload mass allowance. A 3 meter long jointed robot arm is extended from the Dragon capsule's crew hatch, retrieves the sample tray and inserts it into the sample canister payload located on the ERV stage. The robot arm has capacity to obtain grab samples in the event of a rover failure. The sample canister has a robot arm capture casting to enable capture by crewed or robot spacecraft when it returns to Earth orbit

Space Logistics: Launch Capabilities

NASA Technical Reports Server (NTRS)

Furnas, Randall B.

1989-01-01

The current maximum launch capability for the United States are shown. The predicted Earth-to-orbit requirements for the United States are presented. Contrasting the two indicates the strong National need for a major increase in Earth-to-orbit lift capability. Approximate weights for planned payloads are shown. NASA is studying the following options to meet the need for a new heavy-lift capability by mid to late 1990's: (1) Shuttle-C for near term (include growth versions); and (2) the Advanced Lauching System (ALS) for the long term. The current baseline two-engine Shuttle-C has a 15 x 82 ft payload bay and an expected lift capability of 82,000 lb to Low Earth Orbit. Several options are being considered which have expanded diameter payload bays. A three-engine Shuttle-C with an expected lift of 145,000 lb to LEO is being evaluated as well. The Advanced Launch System (ALS) is a potential joint development between the Air Force and NASA. This program is focused toward long-term launch requirements, specifically beyond the year 2000. The basic approach is to develop a family of vehicles with the same high reliability as the Shuttle system, yet offering a much greater lift capability at a greatly reduced cost (per pound of payload). The ALS unmanned family of vehicles will provide a low end lift capability equivalent to Titan IV, and a high end lift capability greater than the Soviet Energia if requirements for such a high-end vehicle are defined.In conclusion, the planning of the next generation space telescope should not be constrained to the current launch vehicles. New vehicle designs will be driven by the needs of anticipated heavy users.

Flight Test of a 40-Foot Nominal Diameter Disk-Gap-Band Parachute Deployed at a Mach Number of 2.72 and a Dynamic Pressure of 9.7 Pounds per Square Foot

NASA Technical Reports Server (NTRS)

Eckstrom, Clinton V.; Preisser, John S.

1968-01-01

A 40-foot-nominal-diameter (12.2 meter) disk-gap-band parachute was flight tested as part of the NASA Supersonic Planetary Entry Decelerator (SPED-I) Program. The test parachute was deployed from an instrumented payload by means of a deployment mortar when the payload was at an altitude of 158,500 feet (48.2 kilometers), a Mach number of 2.72, and a free-stream dynamic pressure of 9.7 pounds per foot(exp 2) (465 newtons per meter(exp 2)). Suspension line stretch occurred 0.46 second after mortar firing and the resulting snatch force loading was -8.lg. The maximum acceleration experienced by the payload due to parachute opening was -27.2g at 0.50 second after the snatch force peak for a total elapsed time from mortar firing of 0.96 second. Canopy-shape variations occurred during the higher Mach number portion of the flight test (M greater than 1.4) and the payload was subjected to large amplitude oscillatory loads. A calculated average nominal axial-force coefficient ranged from about 0.25 immediately after the first canopy opening to about 0.50 as the canopy attained a steady inflated shape. One gore of the test parachute was damaged when the deployment bag with mortar lid passed through it from behind approximately 2 seconds after deployment was initiated. Although the canopy damage caused by the deployment bag penetration had no apparent effect on the functional capability of the test parachute, it may have affected parachute performance since the average effective drag coefficient of 0.48 was 9 percent less than that of a previously tested parachute of the same configuration.

Thermal Analysis of Iodine Satellite (iSAT)

NASA Technical Reports Server (NTRS)

Mauro, Stephanie

2015-01-01

This paper presents the progress of the thermal analysis and design of the Iodine Satellite (iSAT). The purpose of the iSAT spacecraft (SC) is to demonstrate the ability of the iodine Hall Thruster propulsion system throughout a one year mission in an effort to mature the system for use on future satellites. The benefit of this propulsion system is that it uses a propellant, iodine, that is easy to store and provides a high thrust-to-mass ratio. The spacecraft will also act as a bus for an earth observation payload, the Long Wave Infrared (LWIR) Camera. Four phases of the mission, determined to either be critical to achieving requirements or phases of thermal concern, are modeled. The phases are the Right Ascension of the Ascending Node (RAAN) Change, Altitude Reduction, De-Orbit, and Science Phases. Each phase was modeled in a worst case hot environment and the coldest phase, the Science Phase, was also modeled in a worst case cold environment. The thermal environments of the spacecraft are especially important to model because iSAT has a very high power density. The satellite is the size of a 12 unit cubesat, and dissipates slightly more than 75 Watts of power as heat at times. The maximum temperatures for several components are above their maximum operational limit for one or more cases. The analysis done for the first Design and Analysis Cycle (DAC1) showed that many components were above or within 5 degrees Centigrade of their maximum operation limit. The battery is a component of concern because although it is not over its operational temperature limit, efficiency greatly decreases if it operates at the currently predicted temperatures. In the second Design and Analysis Cycle (DAC2), many steps were taken to mitigate the overheating of components, including isolating several high temperature components, removal of components, and rearrangement of systems. These changes have greatly increased the thermal margin available.

Ares V: Enabling Unprecedented Payloads for Space in the 21st Century

NASA Technical Reports Server (NTRS)

Creech, Steve

2010-01-01

Numerous technical and programmatic studies since the U.S. space program began in the 1960s has emphasized the need for a heavy lift capability for exploration beyond low Earth orbit (LEO). The Saturn V once embodied that capability until it was retired. Now the Ares V cargo launch vehicle (CaLV) promises to restore and improve on that capability, providing unprecedented opportunities for human and robotic exploration, science, national security and commercial uses. This paper provides an overview of the capabilities of Ares V, both as an opportunity for payloads of increased mass and/or volume, and as a means of reducing risk in the payload design process. The Ares V is part of NASA s Constellation Program, which also includes the Ares I crew launch vehicle (CLV), Orion crew exploration vehicle (CEV), and Altair lunar lander. This architecture is designed to carry out the national space policy goals of completing the International Space Station (ISS), retiring the Space Shuttle fleet, and expanding human exploration beyond LEO. The Ares V is designed to loft upper stages and/or cargo, such as the Altair lander, into LEO. The Ares I is designed to put Orion into LEO with a crew of up to four for rendezvous with the ISS or with the Ares V Earth departure stage for journeys to the Moon. While retaining the goals of heritage hardware and commonality, the Ares V configuration continues to be refined through a series of internal trades. The current reference configuration was recommended by the Ares Projects and approved by the Constellation Program during the Lunar Capabilities Concept Review (LCCR) June 2008. The reference configuration defines the Ares V as 381 feet (116m) tall with a gross lift-off mass (GLOM) of 8.1 million pounds (3,704.5 mT). Its first stage will generate 11 million pounds of sea-level liftoff thrust. It will be capable of launching 413,800 pounds (187.7 mT) to LEO, 138,500 pounds (63 mT) direct to the Moon or 156,700 pounds (71.1 mT) in its dual-launch architecture role with Ares I. It could also launch 123,100 pounds (55.8 mT) to Sun-Earth L2. Assessment of astronomy payload requirements since Spring 2008 has indicated that Ares V has the potential to support a range of payloads and missions. Some of these missions were impossible in the absence of Ares V s capabilities. Collaborative design/architecture inputs, exchanges, and analyses have already begun between scientists and payload developers. A 2008 study by a National Research Council (NRC) panel, as well as analyses presented by astronomers and planetary scientists at two weekend conferences in 2008, support the position that Ares V has benefit to a broad range of planetary and astronomy missions. This early dialogue with Ares V engineers is permitting the greatest opportunity for payload/transportation/mission synergy and with the least financial impact to Ares V development. In addition, independent analyses suggest that Ares V has the opportunity to enable more cost-effective mission design. 1

Orienting Arc Lamps for Longest Life

NASA Technical Reports Server (NTRS)

Kiss, J.

1985-01-01

Temperature distribution strongly affects performance. Tests on floodlights for Space Shuttle payload bay show useful life of metal halide dc arc lamp prolonged by mounting "anode down" and wiring for maximum heat conduction away from electrodes. Anode-down configuration, anode and cathode temperatures stabilize at 333 degrees and 313 degrees C, respectively, after 1 hour of operation. Temperatures both below limit for quartz-to-metal seals, and lamps able to withstand a 2,000-hour life test with satisfactory light output at end.

First assembly times and equilibration in stochastic coagulation-fragmentation

DOE Office of Scientific and Technical Information (OSTI.GOV)

D’Orsogna, Maria R.; Department of Mathematics, CSUN, Los Angeles, California 91330-8313; Lei, Qi

2015-07-07

We develop a fully stochastic theory for coagulation and fragmentation (CF) in a finite system with a maximum cluster size constraint. The process is modeled using a high-dimensional master equation for the probabilities of cluster configurations. For certain realizations of total mass and maximum cluster sizes, we find exact analytical results for the expected equilibrium cluster distributions. If coagulation is fast relative to fragmentation and if the total system mass is indivisible by the mass of the largest allowed cluster, we find a mean cluster-size distribution that is strikingly broader than that predicted by the corresponding mass-action equations. Combinations ofmore » total mass and maximum cluster size under which equilibration is accelerated, eluding late-stage coarsening, are also delineated. Finally, we compute the mean time it takes particles to first assemble into a maximum-sized cluster. Through careful state-space enumeration, the scaling of mean assembly times is derived for all combinations of total mass and maximum cluster size. We find that CF accelerates assembly relative to monomer kinetic only in special cases. All of our results hold in the infinite system limit and can be only derived from a high-dimensional discrete stochastic model, highlighting how classical mass-action models of self-assembly can fail.« less

Risks of nuclear waste disposal in space. III - Long-term orbital evolution of small particle distribution

NASA Technical Reports Server (NTRS)

Friedlander, A. L.; Wells, W. C.

1980-01-01

A study of long term risks is presented that treats an additional pathway that could result in earth reentry, namely, small radioactive particles released in solar orbit due to payload fragmentation by accidental explosion or meteoroid impact. A characterization of such an event and of the initial mass size distribution of particles is given for two extremes of waste form strength. Attention is given to numerical results showing the mass-time distribution of material and the fraction of initial mass intercepted by earth. It is concluded that it appears that program planners need not be to concerned about the risks of this particular failure mechanism and return pathway.

Miniaturized GC/MS instrumentation for in situ measurements: micro gas chromatography coupled with miniature quadrupole array and paul ion trap mass spectrometers

NASA Technical Reports Server (NTRS)

Holland, P.; Chutjian, A.; Darrach, M.; Orient, O.

2002-01-01

Miniaturized chemical instrumentation is needed for in situ measurements in planetary exploration and other spaceflight applications where factors such as reduction in payload requirements and enhanced robustness are important. In response to this need, we are 'continuing to develop miniaturized GC/MS instrumentation which combines chemical separations by gas chromatography (GC) with mass spectrometry (MS) to provide positive identification of chemical compounds in complex mixtures of gases, such as those found in the International Space Station's cabin atmosphere. Our design approach utilizes micro gas chromatography components coupled with either a miniature quadrupole mass spectrometer array (QMSA) or compact, high-resolution Paul ion trap.

The Near Earth Object Scout Spacecraft: A Low Cost Approach to in-situ Characterization of the NEO Population

NASA Technical Reports Server (NTRS)

Koontz, Steven L.; Condon, Gerald; Graham, Lee; Bevilacqua, Ricardo

2014-01-01

In this paper we describe a micro/nano satellite spacecraft and a supporting mission profile and architecture designed to enable preliminary in-situ characterization of a significant number of Near Earth Objects (NEOs) at reasonable cost. The spacecraft will be referred to as the NEO Scout. NEO Scout spacecraft are to be placed in GTO, GEO, or cis-lunar space as secondary payloads on launch vehicles headed for GTO or beyond and will begin their mission after deployment from the launcher. A distinguishing key feature of the NEO scout system is to design the mission timeline and spacecraft to rendezvous with and land on the target NEOs during close approach to the Earth-Moon system using low-thrust/high- impulse propulsion systems. Mission feasibility and preliminary design analysis are presented along with detailed trajectory calculations. The use of micro/nano satellites in low-cost interplanetary exploration is attracting increasing attention and is the subject of several annual workshops and published design studies (1-4). The NEO population consists of those asteroids and short period comets orbiting the Sun with a perihelion of 1.3 astronomical units or less (5-8). As of July 30, 2013 10065 Near-Earth objects have been discovered. The spin rate, mass, density, surface physical (especially mechanical) properties, composition, and mineralogy of the vast majority of these objects are highly uncertain and the limited available telescopic remote sensing data imply a very diverse population (5-8). In-situ measurements by robotic spacecraft are urgently needed to provide the characterization data needed to support hardware and mission design for more ambitious human and robotic NEO operations. Large numbers of NEOs move into close proximity with the Earth-Moon system every year (9). The JPL Near-Earth Object Human Space Flight Accessible Targets Study (NHATS) (10) has produced detailed mission profile and delta V requirements for various NEO missions ranging from 30 to 420 days in duration and assuming chemical propulsion. Similar studies have been reported assuming high power electric propulsion for manned NEO rendezvous missions (11). The delta V requirement breakdown and mission profile data from references 10 and 11 are used as a basis for sizing the NEO Scout spacecraft and for conducting preliminary feasibility assessments using the Tsiokolvsky rocket equation, a (worst-case) delta V requirement of 10 km/sec, and a maximum spacecraft dry mass of 20 kg. Using chemical propellant for a 10 km/sec delta V drives spacecraft wet mass well above 300 kg so that chemical propulsion is a non-starter for the proposed mission profile and spacecraft wet mass limits. In contrast, a solar electric propulsion system needs only 8 kg of Xe propellant to accelerate the spacecraft to 10 km/sec in 163 days with 0.02 N of thrust and 500 W of power from1.6 sq m of 29% efficient solar panels. In a second example, accelerating a 4 kg payload to 7 km/sec over 180 days requires about 6.7 kg of propellant and 1.2 kg of solar panels (12 kg total spacecraft wet mass).

KSC-00pp1681

NASA Image and Video Library

2000-11-10

In the Space Station Processing Facility, workers attach an overhead crane to lift the P6 integrated truss segment from a workstand and move it to the payload transport canister for transfer to Launch Pad 39B. There it will be placed in Endeavour’s payload bay for launch on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the International Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch is scheduled for Nov. 30 at 10:06 p.m. EST

KSC-00pp1685

NASA Image and Video Library

2000-11-10

As it travels across the Space Station Processing Facility, the P6 integrated truss segment passes over the two Italian-built Multi-Purpose Logistics Modules, Leonardo (right) and Raffaello (behind Leonardo). The P6 is being moved to a payload transport canister for transfer to Launch Pad 39B. There it will be placed in Endeavour’s payload bay for launch on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1684

NASA Image and Video Library

2000-11-10

In the Space Station Processing Facility, the P6 integrated truss segment travels across the building to a payload transport canister for transfer to Launch Pad 39B. There it will be placed in Endeavour’s payload bay for launch on mission STS-97. At left is the airlock module, another component of the International Space Station. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. Launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1686

NASA Image and Video Library

2000-11-10

Carried by an overhead crane, the P6 integrated truss segment travels the length of the Space Station Processing Facility toward a payload transport canister that will transfer it to Launch Pad 39B. At the pad, the Space Station element will be placed in Endeavour’s payload bay for launch on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

KSC-00pp1687

NASA Image and Video Library

2000-11-10

The P6 integrated truss segment hangs suspended from an overhead crane that is moving it the length of the Space Station Processing Facility toward a payload transport canister for transfer to Launch Pad 39B. At the pad, the Space Station element will be placed in Endeavour’s payload bay for launch on mission STS-97. The P6 comprises Solar Array Wing-3 and the Integrated Electronic Assembly, to be installed on the Space Station. The Station’s electrical power system will use eight photovoltaic solar arrays, each 112 feet long by 39 feet wide, to convert sunlight to electricity. The solar arrays are mounted on a “blanket” that can be folded like an accordion for delivery. Once in orbit, astronauts will deploy the blankets to their full size. Gimbals will be used to rotate the arrays so that they will face the Sun to provide maximum power to the Space Station. The STS-97 launch is scheduled Nov. 30 at 10:06 p.m. EST

Astrometric Telescope Facility preliminary systems definition study. Volume 1: Executive summary

NASA Technical Reports Server (NTRS)

Sobeck, Charlie

1987-01-01

The Astrometric Telescope Facility (ATF) is a spaceborne observatory proposed for use on the Space Station (SS) as an Initial Operating Capability (IOC) payload. The primary objective of the ATF will be the search for extrasolar planetary systems and a detailed investigation of any discovered systems. In addition, it will have the capability of conducting other astrophysics investigations; e.g., measuring precise distances and motions of stars within our galaxy. The purposes of the study were to: (1) define mission and system requirements; (2) define a strawman system concept for the facility at the Prephase A level; (3) define the need for additional trade studies or technology development; and (4) estimate program cost for the strawman concept. It has been assumed for the study that the ATF will be a SS payload, will use a SS-provided Coarse Pointing System (CPS), will meet SS constraints, and will make maximum use of existing flight qualified designs or designs to be qualified by the SS program for general SS use.

An analysis of the effect of aeroassist maneuvers on orbital transfer vehicle performance

NASA Technical Reports Server (NTRS)

Murphy, Gregory O.; Suit, William T.

1987-01-01

This paper summarizes a Langley Research Summer Scholars (LARSS) research project (Summer 1986) dealing with the topic of the effectiveness of aeroassist maneuvers to accomplish a change in the orbital inclination of an Orbital Transfer Vehicle (OTV). This task was subject to OTV design constraints, chief of which were the axial acceleration and the aerodynamic heating rate limits of the OTV. The use of vehicle thrust to replace lost kinetic energy and, thereby, to increase the maximum possible change in orbital inclination was investigated. A relation between time in the hover orbit and payload to LEO was established. The amount of plane change possible during this type of maneuver was checked for several runs and a possible thrusting procedure to increase the plane change and still get to LEO was suggested. Finally, the sensitivity of various target parameters to controllable independent variables was established, trades between the amount of control allowed, and payload to LEO suggested.

NASA Affordable Vehicle Avionics (AVA). Common Modular Avionics System for Nanolaunchers Offering Affordable Access to Space; [Space Technology: Game Changing Development

NASA Technical Reports Server (NTRS)

Aquilina, Rudy

2017-01-01

Small satellites are becoming ever more capable of performing valuable missions for both government and commercial customers. However, currently these satellites can be launched affordably only as secondary payloads. This makes it difficult for the small satellite mission to launch when needed, to the desired orbit, and with acceptable risk. What is needed is a class of low-cost launchers, so that launch costs to low-Earth orbit (LEO) are commensurate with payload costs. Several private and government-sponsored launch vehicle developers are working toward just that-the ability to affordably insert small payloads into LEO. But until now, cost of the complex avionics remained disproportionately high. AVA (Affordable Vehicle Avionics) solves this problem. Significant contributors to the cost of launching nanosatellites to orbit are the avionics and software systems that steer and control the launch vehicles, sequence stage separation, deploy payloads, and telemeter data. The high costs of these guidance, navigation and control (GNC) avionics systems are due in part to the current practice of developing unique, single-use hardware and software for each launch. High-performance, high-reliability inertial sensors components with heritage from legacy launchers also contribute to costs-but can low-cost commercial inertial sensors work just as well? NASA Ames Research Center has developed and tested a prototype low-cost avionics package for space launch vehicles that provides complete GNC functionality in a package smaller than a tissue box (100 millimeters by 120 millimeters by 69 millimeters; 4 inches by 4.7 inches by 2.7 inches), with a mass of less than 0.84 kilogram (2 pounds. AVA takes advantage of commercially available, low-cost, mass-produced, miniaturized sensors, filtering their more noisy inertial data with real-time GPS (Global Positioning Satellite) data. The goal of the AVA project is to produce and light-verify a common suite of avionics and software that deliver affordable, capable GNC and telemetry avionics with application to multiple nanolaunch vehicles at 1 percent of the cost of current state-of-the-art avionics.

Astronomical reach of fundamental physics.

PubMed

Burrows, Adam S; Ostriker, Jeremiah P

2014-02-18

Using basic physical arguments, we derive by dimensional and physical analysis the characteristic masses and sizes of important objects in the universe in terms of just a few fundamental constants. This exercise illustrates the unifying power of physics and the profound connections between the small and the large in the cosmos we inhabit. We focus on the minimum and maximum masses of normal stars, the corresponding quantities for neutron stars, the maximum mass of a rocky planet, the maximum mass of a white dwarf, and the mass of a typical galaxy. To zeroth order, we show that all these masses can be expressed in terms of either the Planck mass or the Chandrasekar mass, in combination with various dimensionless quantities. With these examples, we expose the deep interrelationships imposed by nature between disparate realms of the universe and the amazing consequences of the unifying character of physical law.

Astronomical reach of fundamental physics

PubMed Central

Burrows, Adam S.; Ostriker, Jeremiah P.

2014-01-01

Using basic physical arguments, we derive by dimensional and physical analysis the characteristic masses and sizes of important objects in the universe in terms of just a few fundamental constants. This exercise illustrates the unifying power of physics and the profound connections between the small and the large in the cosmos we inhabit. We focus on the minimum and maximum masses of normal stars, the corresponding quantities for neutron stars, the maximum mass of a rocky planet, the maximum mass of a white dwarf, and the mass of a typical galaxy. To zeroth order, we show that all these masses can be expressed in terms of either the Planck mass or the Chandrasekar mass, in combination with various dimensionless quantities. With these examples, we expose the deep interrelationships imposed by nature between disparate realms of the universe and the amazing consequences of the unifying character of physical law. PMID:24477692

Relation of maximum blood pressure during exercise and regular physical activity in normotensive men with left ventricular mass and hypertrophy. MARATHOM Investigators. Medida de la Actividad fisica y su Relación Ambiental con Todos los Lípidos en el HOMbre.

PubMed

Molina, L; Elosua, R; Marrugat, J; Pons, S

1999-10-15

The relation between maximum systolic blood pressure (BP) during exercise and left ventricular (LV) mass is controversial. Physical activity also induces LV mass increase. The objective was to assess the relation between BP response to exercise and LV mass in normotensive men, taking into account physical activity practice. A cross-sectional study was performed. Three hundred eighteen healthy normotensive men, aged between 20 and 60 years, participated in this study. The Minnesota questionnaire was used to assess physical activity practice. An echocardiogram and a maximum exercise test were performed. LV mass was calculated and indexed to body surface area. LV hypertrophy was defined as a ventricular mass index > or =134 g/m2. BP was measured at the moment of maximum effort. Hypertensive response was considered when BP was > or =210 mm Hg. In the multiple linear regression model, maximum systolic BP was associated with LV mass index and correlation coefficient was 0.27 (SE 0.07). Physical activity practice and age were also associated with LV mass. An association between hypertensive response to exercise and LV hypertrophy was observed (odds ratio 3.16). Thus, BP response to exercise is associated with LV mass and men with systolic BP response > or =210 mm Hg present a 3-times higher risk of LV hypertrophy than those not reaching this limit. Physical activity practice is related to LV mass, but not to LV hypertrophy.

Improving of Hybrid Rocket Engine on the Basis of Optimizing Design Fuel Grain

NASA Astrophysics Data System (ADS)

Oriekov, K. M.; Ushkin, M. P.

2015-09-01

This article examines the processes intrachamber in hybrid rocket engine (HRE) and the comparative assessment of the use of solid rocket motors (SRM) and HRE for meteorological rockets with a mass of payload of the 364 kg. Results of the research showed the possibility of a significant increase in the ballistic effectiveness of meteorological rocket.

Effect of Link Flexibility on tip position of a single link robotic arm

NASA Astrophysics Data System (ADS)

Madhusudan Raju, E.; Siva Rama Krishna, L.; Mouli, Y. Sharath Chandra; Nageswara Rao, V.

2015-12-01

The flexible robots are widely used in space applications due to their quick response, lower energy consumption, lower overall mass and operation at high speed compared to conventional industrial rigid link robots. These robots are inherently flexible, so that the kinematics of flexible robots can't be solved with rigid body assumptions. The flexibility in links and joints affects end-point positioning accuracy of the robot. It is important to model the link kinematics with precision which in turn simplifies modelling of dynamics of flexible robots. The main objective of this paper is to evaluate the effect of link flexibility on a tip position of a single link robotic arm for a given motion. The joint is assumed to be rigid and only link flexibility is considered. The kinematics of flexible link problem is evaluated by Assumed Modes Method (AMM) using MAT LAB Programming. To evaluate the effect of link flexibility (with and without payload) of robotic arm, the normalized tip deviation is found for flexible link with respect to a rigid link. Finally, the limiting inertia for payload mass is found if the allowable tip deviation is 5%.

Optical PAyload for Lasercomm Science (OPALS) link validation

NASA Technical Reports Server (NTRS)

Biswas, Abhijit; Oaida, Bogdan V.; Andrews, Kenneth S.; Kovalik, Joseph M.; Abrahamson, Matthew J.; Wright, Malcolm W.

2015-01-01

Recently several day and nighttime links under diverse atmospheric conditions were completed using the Optical Payload for Lasercomm Science (OPALS) flight system on-board the International Space Station (ISS). In this paper we compare measured optical power and its variance at either end of the link with predictions that include atmospheric propagation models. For the 976 nm laser beacon mean power transmitted from the ground to the ISS the predicted mean irradiance of 10's of microwatts per square meter close to zenith and its decrease with range and increased air mass shows good agreement with predictions. The irradiance fluctuations sampled at 100 Hz also follow the expected increase in scintillation with air mass representative of atmospheric coherence lengths at zenith at 500 nm in the 3-8 cm range. The downlink predicted power of 100's of nanowatts was also reconciled within the uncertainty of the atmospheric losses. Expected link performance with uncoded bit-error rates less than 1E-4 required for the Reed-Solomon code to correct errors for video, text and file transmission was verified. The results of predicted and measured powers and fluctuations suggest the need for further study and refinement.

Evaluation of inertial devices for the control of large, flexible, space-based telerobotic arms

NASA Technical Reports Server (NTRS)

Montgomery, Raymond C.; Kenny, Sean P.; Ghosh, Dave; Shenhar, Joram

1993-01-01

Inertial devices, including sensors and actuators, offer the potential of improving the tracking of telerobotic commands for space-based robots by smoothing payload motions and suppressing vibrations. In this paper, inertial actuators (specifically, torque-wheels and reaction-masses) are studied for that potential application. Batch simulation studies are presented which show that torque-wheels can reduce the overshoot in abrupt stop commands by 82 percent for a two-link arm. For man-in-the-loop evaluation, a real-time simulator has been developed which samples a hand-controller, solves the nonlinear equations of motion, and graphically displays the resulting motion on a computer workstation. Currently, two manipulator models, a two-link, rigid arm and a single-link, flexible arm, have been studied. Results are presented which show that, for a single-link arm, a reaction-mass/torque-wheel combination at the payload end can yield a settling time of 3 s for disturbances in the first flexible mode as opposed to 10 s using only a hub motor. A hardware apparatus, which consists of a single-link, highly flexible arm with a hub motor and a torque-wheel, has been assembled to evaluate the concept and is described herein.

Multiplying Mars Lander Opportunities with Marsdrop Microlanders

NASA Technical Reports Server (NTRS)

Staehle, Robert L.; Spangelo, Sara; Lane, Marc S.; Aaron, Kim M.; Bhartia, Rohit; Boland, Justin S.; Christensen, Lance E.; Forouhar, Siamak; de la Torre Juarez, Manuel; Trawny, Nikolas;

Evaluation of inertial devices for the control of large, flexible, space-based telerobotic arms

NASA Astrophysics Data System (ADS)

Montgomery, Raymond C.; Kenny, Sean P.; Ghosh, Dave; Shenhar, Joram

1993-02-01

Inertial devices, including sensors and actuators, offer the potential of improving the tracking of telerobotic commands for space-based robots by smoothing payload motions and suppressing vibrations. In this paper, inertial actuators (specifically, torque-wheels and reaction-masses) are studied for that potential application. Batch simulation studies are presented which show that torque-wheels can reduce the overshoot in abrupt stop commands by 82 percent for a two-link arm. For man-in-the-loop evaluation, a real-time simulator has been developed which samples a hand-controller, solves the nonlinear equations of motion, and graphically displays the resulting motion on a computer workstation. Currently, two manipulator models, a two-link, rigid arm and a single-link, flexible arm, have been studied. Results are presented which show that, for a single-link arm, a reaction-mass/torque-wheel combination at the payload end can yield a settling time of 3 s for disturbances in the first flexible mode as opposed to 10 s using only a hub motor. A hardware apparatus, which consists of a single-link, highly flexible arm with a hub motor and a torque-wheel, has been assembled to evaluate the concept and is described herein.

Conformal Ablative Thermal Protection System for Planetary and Human Exploration Missions: Overview of the Technology Maturation Efforts Funded by NASA's Game Changing Development Program

NASA Technical Reports Server (NTRS)

Beck, Robin A.; Arnold, James O.; Gasch, Matthew J.; Stackpoole, Margaret M.; Fan, Wendy; Szalai, Christine E.; Wercinski, Paul F.; Venkatapathy, Ethiraj

2012-01-01

The Office of Chief Technologist (OCT), NASA has identified the need for research and technology development in part from NASA's Strategic Goal 3.3 of the NASA Strategic Plan to develop and demonstrate the critical technologies that will make NASA's exploration, science, and discovery missions more affordable and more capable. Furthermore, the Game Changing Development Program (GCDP) is a primary avenue to achieve the Agency's 2011 strategic goal to "Create the innovative new space technologies for our exploration, science, and economic future." In addition, recently released "NASA space Technology Roadmaps and Priorities," by the National Research Council (NRC) of the National Academy of Sciences stresses the need for NASA to invest in the very near term in specific EDL technologies. The report points out the following challenges (Page 2-38 of the pre-publication copy released on February 1, 2012): Mass to Surface: Develop the ability to deliver more payload to the destination. NASA's future missions will require ever-greater mass delivery capability in order to place scientifically significant instrument packages on distant bodies of interest, to facilitate sample returns from bodies of interest, and to enable human exploration of planets such as Mars. As the maximum mass that can be delivered to an entry interface is fixed for a given launch system and trajectory design, the mass delivered to the surface will require reduction in spacecraft structural mass; more efficient, lighter thermal protection systems; more efficient lighter propulsion systems; and lighter, more efficient deceleration systems. Surface Access: Increase the ability to land at a variety of planetary locales and at a variety of times. Access to specific sites can be achieved via landing at a specific location (s) or transit from a single designated landing location, but it is currently infeasible to transit long distances and through extremely rugged terrain, requiring landing close to the site of interest. The entry environment is not always guaranteed with a direct entry, and improving the entry system's robustness to a variety of environmental conditions could aid in reaching more varied landing sites."

In-flight calibration and performance evaluation of the fixed head star trackers for the solar maximum mission

NASA Technical Reports Server (NTRS)

Thompson, R. H.; Gambardella, P. J.

1980-01-01

The Solar Maximum Mission (SMM) spacecraft provides an excellent opportunity for evaluating attitude determination accuracies achievable with tracking instruments such as fixed head star trackers (FHSTs). As a part of its payload, SMM carries a highly accurate fine pointing Sun sensor (FPSS). The EPSS provides an independent check of the pitch and yaw parameters computed from observations of stars in the FHST field of view. A method to determine the alignment of the FHSTs relative to the FPSS using spacecraft data is applied. Two methods that were used to determine distortions in the 8 degree by 8 degree field of view of the FHSTs using spacecraft data are also presented. The attitude determination accuracy performance of the in flight calibrated FHSTs is evaluated.

Non-rocket Earth-Moon transportation system

NASA Astrophysics Data System (ADS)

Bolonkin, A.

Author suggests and researches one of his methods of flights to outer Space, described in book "Non Rocket Flights in Space", which is prepared and offered for publication. In given report the method and facilities named "Bolonkin Transport System" (BTS) for delivering of payload and people to Moon and back is presented. BTS can be used also for free trip to outer Space up at altitude 60,000 km and more. BTS can be applying as a trust system for atmospheric supersonic aircrafts, and as a free energy source. This method uses, in general, the rotary and kinetic energy of the Moon. The manuscript contains the theory and results of computation of special Project. This project uses three cables (main and two for driving of loads) from artificial material: fiber, whiskers, nanotubes, with the specific tensile strength (ratio the tensile stress to density) k=/=4*10^7 or more. The nanotubes with same and better parameters are received in scientific laboratories. Theoretical limit of nanotubes SWNT is about k=100*10^7. The upper end of the cable is connected to the Moon. The lower end of the cable is connected to an aircraft (or buoy), which flies (i.e. glides or slides) in Earth atmosphere along the planet's surface. The aircraft (and Moon) has devices, which allows the length of cables to be changed. The device would consists of a spool, motor, brake, transmission, and controller. The facility could have devices for delivering people and payloads t o the Moon and back using the suggested Transport System. The delivery devices include: containers, cables, motors, brakes, and controllers. If the aircraft is small and the cable is strong the motion of the Moon can be used to move the airplane. For example (see enclosed project), if the airplane weighs 15 tons and has an aerodynamic ratio (the lift force to the drag force) equal 5, a thrust of 3000 kg would be enough for the aircraft to fly for infinity without requiring any fuel. The aircraft could use a small turbine engine for maneuverability and temporary landing on Earth surface up to 6 hours. If the suggested Transport System is used only for free thrust (9 tons), the system can pull the three supersonic transport (passenger) aircraft or produces up to 40 millions watt energy. The Moon's trajectory has an eccentricity. If the main cable is strong enough, the Moon may be used for free lifting (traveling) a payload (men cabins) to an altitude of about 60,000 kilometers every 27 days. For this case, the length of the main cable from the Moon to the container does not change and when the Moon increases its distance from the Moon to the Earth, the Moon lifts the space ship. The cabin could land back on the planet at any time if it is allowed to slide along the cable. The Moon's energy can be used also for an inexpensive trip (passenger system) around of the Earth by having the moon "drag" an aircraft around the planet (using the Moon as free thrust engine) with supersonic speed, about 440 m/s (Max Number is 1.5). Project . Free Trips to the Moon. The following is some data estimating the Moon Transport System which provides an inexpensive payload transfer between the Earth and the Moon. The system has three cables. Every cable can keep the force 3 tons. Material of cable has specific strong k=4*10^7. All cables have cross-sectional areas of an equal stress. A minimal cross-sectional area of cable is 0.42 mm2 (diameter 0,73 mm) and maximum cross-sectional area is 1.9 mm2 (1.56 mm). The mass of the main cable is 1300 tons. The total mass of the main cable plus the two container cables (for delivering a mass 3000 kg) equals 3900 tons. An unlimited and inexpensive means of the payload delivery between the Earth and the Moon could be obtained. An elapsed time to the Moon trip with a speed of 6 km/sec would be about 18.5 hours. The annual deliver capably is 1320 tons (or 10,000 tourists) in every direction. If one cable will be damaged two others will be used for safe people and repairing. The author has also solved additional problems, which appear in these projects and which can look as difficult for the proposed space transportation technology. The author is prepared to discuss the problems with serious organizations, which want to research and develop these projects. Patent application is 09/789,959 USA. See also other my proposals to WSC 2002.

A Mobile Asteroid Surface Scout for the AIDA Mission

NASA Astrophysics Data System (ADS)

Ho, Tra Mi; Lange, Caroline; Grimm, Christian; Thimo Grundmann, Jan; Rößler, Johannes; Schröder, Silvio; Skoczylas, Thomas; Ziach, Christian; Biele, Jens; Cozzoni, Barbara; Krause, Christian; Küchemann, Oliver; Maibaum, Michael; Ulamec, Stephan; Lange, Michael; Mierheim, Olaf; Maier, Maximilian; Herique, Alain; Mascot Study Team

2016-04-01

The Asteroid Impact Deflection, AIDA, mission is composed of a kinetic impactor, DART and an observer, the Asteroid Impact Monitor, AIM, carrying among other payload a surface package, MASCOT2 (MSC2). Its proposed concept is based on the MASCOT lander onboard the HAYABUSA2 Mission (JAXA) to near-Earth asteroid (162173) Ryugu. MASCOT is a compact platform ('shoe box size') carrying a suite of 4 scientific instruments and has a landed mass of ~10kg. Equipped with a mobility mechanism, the MASCOT lander is able to upright and relocate on the targeted asteroid; thus providing in-situ data at more than one site. In the context of the AIDA Mission, the MASCOT2 lander would be carried by the AIM spacecraft and delivered onto Didymoon, the secondary object in the (65803) Didymos binary near-Earth asteroid system. Since the mission objectives of the AIM mission within the joint AIDA mission concept differ from JAXA's sample return mission HAYABUSA2, several design changes need to be studied and implemented. To support one of the prime objectives of the AIM mission, the characterization of the bulk physical properties of Didymoon, the main scientific payload of MSC2 is a low-frequency radar (LFR) to investigate the internal structure of the asteroid moon. Since the total science payload on MASCOT2 is limited to approximately 2.3 kg, the mass remaining for a suite of other experiments is in the range of 0.1 to 0.5 kg per instrument. Further requirements have a significant impact on the MSC2 design which will be presented. Among these are the much longer required operational lifetime than for MASCOT on HAYABUSA2, and different conditions on the target body such as an extremely low gravity due to its small size of Ø_[Didymoon] ~ 150m.

NASA's Space Launch System: Building a New Capability for Discovery

NASA Technical Reports Server (NTRS)

Creech, Stephen D.; Robinson, Kimberly F.

2015-01-01

Designed to enable human space exploration missions, including eventually landings on Mars, NASA's Space Launch System (SLS) represents a unique launch capability with a wide range of utilization opportunities, from delivering habitation systems into the lunar vicinity to high-energy transits through the outer solar system. Substantial progress has been made toward the first launch of the initial configuration of SLS, which will be able to deliver more than 70 metric tons of payload into low Earth orbit (LEO). The vehicle will then be evolved into more powerful configurations, culminating with the capability to deliver more than 130 metric tons to LEO. The initial configuration will be able to deliver greater mass to orbit than any contemporary launch vehicle, and the evolved configuration will have greater performance than the Saturn V rocket that enabled human landings on the moon. SLS will also be able to carry larger payload fairings than any contemporary launch vehicle, and will offer opportunities for co-manifested and secondary payloads. Because of its substantial mass-lift capability, SLS will also offer unrivaled departure energy, enabling mission profiles currently not possible. The basic capabilities of SLS have been driven by studies on the requirements of human deep-space exploration missions, and continue to be validated by maturing analysis of Mars mission options. Early collaboration with science teams planning future decadal-class missions have contributed to a greater understanding of the vehicle's potential range of utilization. As this paper will explain, SLS is making measurable progress toward becoming a global infrastructure asset for robotic and human scouts of all nations by providing the robust space launch capability to deliver sustainable solutions for exploration.

NASA'S Space Launch System Mission Capabilities for Exploration

NASA Technical Reports Server (NTRS)

Creech, Stephen D.; Crumbly, Christopher M.; Robinson, Kimberly F.

2015-01-01

Designed to enable human space exploration missions, including eventual landings on Mars, NASA’s Space Launch System (SLS) represents a unique launch capability with a wide range of utilization opportunities, from delivering habitation systems into the lunar vicinity to high-energy transits through the outer solar system. Developed with the goals of safety, affordability and sustainability in mind, SLS is a foundational capability for NASA’s future plans for exploration, along with the Orion crew vehicle and upgraded ground systems at the agency’s Kennedy Space Center. Substantial progress has been made toward the first launch of the initial configuration of SLS, which will be able to deliver more than 70 metric tons of payload into low Earth orbit (LEO), greater mass-to-orbit capability than any contemporary launch vehicle. The vehicle will then be evolved into more powerful configurations, culminating with the capability to deliver more than 130 metric tons to LEO, greater even than the Saturn V rocket that enabled human landings on the moon. SLS will also be able to carry larger payload fairings than any contemporary launch vehicle, and will offer opportunities for co-manifested and secondary payloads. Because of its substantial mass-lift capability, SLS will also offer unrivaled departure energy, enabling mission profiles currently not possible. Early collaboration with science teams planning future decadal-class missions have contributed to a greater understanding of the vehicle’s potential range of utilization. This presentation will discuss the potential opportunities this vehicle poses for the planetary sciences community, relating the vehicle’s evolution to practical implications for mission capture. As this paper will explain, SLS will be a global launch infrastructure asset, employing sustainable solutions and technological innovations to deliver capabilities for space exploration to power human and robotic systems beyond our Moon and in to deep space.

NASA's Space Launch System Mission Capabilities for Exploration

NASA Technical Reports Server (NTRS)

Creech, Stephen D.; Crumbly, Christopher M.; Robinson, Kimberly F.

2015-01-01

Designed to enable human space exploration missions, including eventual landings on Mars, NASA's Space Launch System (SLS) represents a unique launch capability with a wide range of utilization opportunities, from delivering habitation systems into the lunar vicinity to high-energy transits through the outer solar system. Developed with the goals of safety, affordability and sustainability in mind, SLS is a foundational capability for NASA's future plans for exploration, along with the Orion crew vehicle and upgraded ground systems at the agency's Kennedy Space Center. Substantial progress has been made toward the first launch of the initial configuration of SLS, which will be able to deliver more than 70 metric tons of payload into low Earth orbit (LEO), greater mass-to-orbit capability than any contemporary launch vehicle. The vehicle will then be evolved into more powerful configurations, culminating with the capability to deliver more than 130 metric tons to LEO, greater even than the Saturn V rocket that enabled human landings on the moon. SLS will also be able to carry larger payload fairings than any contemporary launch vehicle, and will offer opportunities for co-manifested and secondary payloads. Because of its substantial mass-lift capability, SLS will also offer unrivaled departure energy, enabling mission profiles currently not possible. Early collaboration with science teams planning future decadal-class missions have contributed to a greater understanding of the vehicle's potential range of utilization. This presentation will discuss the potential opportunities this vehicle poses for the planetary sciences community, relating the vehicle's evolution to practical implications for mission capture. As this paper will explain, SLS will be a global launch infrastructure asset, employing sustainable solutions and technological innovations to deliver capabilities for space exploration to power human and robotic systems beyond our Moon and in to deep space.

Aerocapture Benefits to Future Science Missions

NASA Technical Reports Server (NTRS)

Artis, Gwen; James, Bonnie

2006-01-01

NASA's In-Space Propulsion Technology (ISPT) Program is investing in technologies to revolutionize the robotic exploration of deep space. One of these technologies is Aerocapture, the most promising of the "aeroassist" techniques used to maneuver a space vehicle within an atmosphere, using aerodynamic forces in lieu of propellant. (Other aeroassist techniques include aeroentry and aerobraking.) Aerocapture relies on drag atmospheric drag to decelerate an incoming spacecraft and capture it into orbit. This technique is very attractive since it permits spacecraft to be launched from Earth at higher velocities, providing shorter trip times and saving mass and overall cost on future missions. Recent aerocapture systems analysis studies quantify the benefits of aerocapture to future exploration. The 2002 Titan aerocapture study showed that using aerocapture at Titan instead of conventional propulsive capture results in over twice as much payload delivered to Titan. Aerocapture at Venus results in almost twice the payload delivered to Venus as with aerobraking, and over six times more mass delivered into orbit than all-propulsive capture. Aerocapture at Mars shows significant benefits as the payload sizes increase and as missions become more complex. Recent Neptune aerocapture studies show that aerocapture opens up entirely new classes of missions at Neptune. Current aerocapture technology development is advancing the maturity of each subsystem technology needed for successful implementation of aerocapture on future missions. Recent development has focused on both rigid aeroshell and inflatable aerocapture systems. Rigid aeroshell systems development includes new ablative and non-ablative thermal protection systems, advanced aeroshell performance sensors, lightweight structures and higher temperature adhesives. Inflatable systems such as trailing tethered and clamped "ballutes" and inflatable aeroshells are also under development. Computational tools required to support future aerocapture missions are an integral part of aerocapture development. Tools include engineering reference atmosphere models, guidance and navigation algorithms, aerothermodynamic modeling, and flight simulation.

Specific Space Transportation Costs to GEO - Past, Present and Future

NASA Astrophysics Data System (ADS)

Koelle, Dietrich E.

2002-01-01

The largest share of space missions is going to the Geosynchronous Orbit (GEO); they have the highest commercial importance. The paper first shows the historic trend of specific transportation costs to GEO from 1963 to 2002. It started out with more than 500 000 /kg(2002-value) and has come down to 36 000 /kg. This reduction looks impressive, however, the reason is NOT improved technology or new techniques but solely the growth of GEO payloads`unit mass. The first GEO satellite in 1963 did have a mass of 36 kg mass (BoL) . This has grown to a weight of 1600 kg (average of all GEO satellites) in the year 2000. Mass in GEO after injection is used here instead of GTO mass since the GTO mass depends on the launch site latitude. The specific cost reduction is only due to the "law-of-scale", valid in the whole transportation business: the larger the payload, the lower the specific transportation cost. The paper shows the actual prices of launch services to GTO by the major launch vehicles. Finally the potential GEO transportation costs of future launch systems are evaluated. What is the potential reduction of specific transportation costs if reusable elements are introduced in future systems ? Examples show that cost reductions up to 75 % seem achievable - compared to actual costs - but only with launch systems optimized according to modern principles of cost engineering. 1. 53rd International Astronautical Congress, World Space Congress Houston 2. First Submission 3. Specific Space Transportation Costs to GEO - Past, Present and Future 4. KOELLE, D.E. 5. IAA.1.1 Launch Vehicles' Cost Engineering and Economic Competitiveness 6. D.E. Koelle; A.E. Goldstein 7. One overhead projector and screen 8. Word file attached 9. KOELLE I have approval to attend the Congress. I am not willing to present this paper at the IAC Public Outreach Program.

Body mass and antler development patterns of Rocky Mountain elk (Cervus elaphus nelsoni) in Michigan

USGS Publications Warehouse

Bender, L.C.; Carlson, E.; Schmitt, S.M.; Haufler, J.B.

2003-01-01

We documented mean and maximum body mass, mass accretion patterns and ander development patterns of Rocky Mountain elk in Michigan. Mean body mass of bulls averaged 9-11% heavier, and maximum body mass 23-27% heavier, in Michigan than in other Rocky Mountain elk populations. Mean live body mass of cows averaged 11% heavier in Michigan, but mean eviscerated body mass did not differ. Maximum body mass of cows was 10-24% heavier in Michigan. Body mass peaked at age 7.5 for bulls and 8.5 for cows, similar to other Rocky Mountain elk populations despite the greater body mass achieved in Michigan. Sexual dimorphism in bull and cow body mass increased until peak body mass was attained, whereupon bulls were ???38% heavier than cows. Antler development of bull elk peaked at age 10.5, comparable to other Rocky Mountain elk populations. Relations between antler development and body mass within age classes were highly variable, but generally weak. Greater body mass seen in Michigan, and the peaking of antler development well after body mass in bulls, suggested a phenotypic response to nutritional conditions that allow Rocky Mountain elk in Michigan to maximize the species growth potential.

Systems analysis of Mars solar electric propulsion vehicles

NASA Technical Reports Server (NTRS)

Hickman, J. M.; Curtis, H. B.; Kenny, B. H.; Sefcik, R. J.

1990-01-01

Mission performance, mass, initial power, and cost are determined for solar electric propulsion vehicles across a range of payload masses, reference powers, and mission trajectories. Thick radiation shielding is added to arrays using indium phosphide or III-V multijunction solar cells to reduce the damage incurred through the radiation belts. Special assessments of power management and distribution systems, atmospheric drag, and energy storage are made. It is determined that atmospheric drag is of no great concern and that the energy storage used in countering drag is unnecessary. A scheme to package the arrays, masts, and ion thrusters into a single fairing is presented.