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NASA NTRS2015

NASA Near Earth Network (NEN), Deep Space Network (DSN) and Space Network (SN) Support of CubeSat Communications

There has been a historical trend to increase capability and drive down the Size, Weight and Power (SWAP) of satellites and that trend continues today. Small satellites, including systems conforming to the CubeSat specification, because of their low launch and development costs, are enabling new concepts and capabilities for science investigations across multiple fields of interest to NASA. NASA scientists and engineers across many of NASAs Mission Directorates and Centers are developing exciting CubeSat concepts and welcome potential partnerships for CubeSat endeavors. From a communications and tracking point of view, small satellites including CubeSats are a challenge to coordinate because of existing small spacecraft constraints, such as limited SWAP and attitude control, low power, and the potential for high numbers of operational spacecraft. The NASA Space Communications and Navigation (SCaN) Programs Near Earth Network (NEN), Deep Space Network (DSN) and the Space Network (SN) are customer driven organizations that provide comprehensive communications services for space assets including data transport between a missions orbiting satellite and its Mission Operations Center (MOC). The NASA NEN consists of multiple ground antennas. The SN consists of a constellation of geosynchronous (Earth orbiting) relay satellites, named the Tracking and Data Relay Satellite System (TDRSS). The DSN currently makes available 13 antennas at its three tracking stations located around the world for interplanetary communication. The presentation will analyze how well these space communication networks are positioned to support the emerging small satellite and CubeSat market. Recognizing the potential support, the presentation will review the basic capabilities of the NEN, DSN and SN in the context of small satellites and will present information about NEN, DSN and SN-compatible flight radios and antenna development activities at the Goddard Space Flight Center (GSFC) and across industry. The presentation will review concepts on how the SN multiple access capability could help locate CubeSats and provide a low-latency early warning system. The presentation will also present how the DSN is evolving to maximize use of its assets for interplanetary CubeSats. The critical spectrum-related topics of available and appropriate frequency bands, licensing, and coordination will be reviewed. Other key considerations, such as standardization of radio frequency interfaces and flight and ground communications hardware systems, will be addressed as such standardization may reduce the amount of time and cost required to obtain frequency authorization and perform compatibility and end-to-end testing. Examples of standardization that exist today are the NASA NEN, DSN and SN systems which have published users guides and defined frequency bands for high data rate communication, as well as conformance to CCSDS standards. The workshop session will also seek input from the workshop participants to better understand the needs of small satellite systems and to identify key development activities and operational approaches necessary to enhance communication and navigation support using NASA's NEN, DSN and SN.

Telecommunication↗

NASA NTRS2012

Interplanetary CubeSats: Opening the Solar System to a Broad Community at Lower Cost

Interplanetary CubeSats could enable small, low-cost missions beyond low Earth orbit. This class is defined by mass < ~ 10 kg, cost < $30 M, and durations up to 5 years. Over the coming decade, a stretch of each of six distinct technology areas, creating one overarching architecture, could enable comparatively low-cost Solar System exploration missions with capabilities far beyond those demonstrated in small satellites to date. The six technology areas are: (1) CubeSat electronics and subsystems extended to operate in the interplanetary environment, especially radiation and duration of operation; (2) Optical telecommunications to enable very small, low-power uplink/downlink over interplanetary distances; (3) Solar sail propulsion to enable high !V maneuvering using no propellant; (4) Navigation of the Interplanetary Superhighway to enable multiple destinations over reasonable mission durations using achievable !V; (5) Small, highly capable instrumentation enabling acquisition of high-quality scientific and exploration information; and (6) Onboard storage and processing of raw instrument data and navigation information to enable maximum utility of uplink and downlink telecom capacity, and minimal operations staffing. The NASA Innovative Advanced Concepts (NIAC) program in 2011 selected Interplanetary CubeSats for further investigation, some results of which are reported here for Phase 1.

Staehle, Robert L.↗

NASA NTRS1969

The Small Scientific Satellite

Small Scientific Satellite program, discussing payload for investigating magnetosphere and near interplanetary space

Williams, D. J.↗

NASA NTRSDate not supplied

The Adaptable, Deployable, Entry and Placement Technology (ADEPT) Enabling Advanced Entry, Descent, and Landing Capabilities for SmallSat Missions

The Adaptable, Deployable Entry and Placement Technology (ADEPT) is a mechanically deployable low-ballistic coefficient aeroshell entry system which performs entry, descent, and landing (EDL) functions for a broad range of planetary destinations including Mars, Venus, Titan and Earth. The deployable system allows mission planners to develop an aeroshell design that is stowed like a folded umbrella, and yet prior to the EDL mission segment, transforms into a deployed, near-rigid low ballistic coefficient configuration. The ADEPT architecture is scalable from a sub-1m class to 6m diameter entry vehicles for Venus and and it has been studied at up to 15m – 20m scale for human exploration class missions. There has been growing interest in performing interplanetary missions with small spacecraft as secondary payloads and ADEPT offers small satellite mission designers the ability to consider Aeroassist technologies to expand mission applications. The ADEPT architecture introduces a new approach for entry vehicle design. It deploys a flexible 3D woven carbon fabric as the primary drag surface. It is capable of surviving high aerothermal heating during entry, while maintaining a robust impermeable membrane that provides high drag for vehicle deceleration. The ADEPT project team has advanced this decelerator technology via systems-level testing at the one meter diameter (nano-ADEPT) scale. A subsonic aeroloads test (May 2015) and a pathfinder arc-jet aeroheating test (Sept 2015) and a successful sounding rocket flight (SR-1) launch of a 0.7 meter deployed diameter ADEPT (September 2018). The SR-1 flight experiment demonstrated most of the primary end-to-end mission stages including: launch in a stowed configuration, separation and deployment in zero-g, exo-atmospheric conditions, and passive ballistic re-entry of a 70 degree half-angle cone geometry. The ADEPT SR-1 sounding rocket flight experiment had to address challenges such as stowing within the constrained payload envelope, fabric folding, and packaging concerns.

Paul Wercinski↗

NASA NTRS2014

Preliminary Analysis: Am-241 RHU/TEG Electric Power Source for Nanosatellites

The Februay 2013 Space Works Commercial report indicates a strong increase in nano/microsatellite (1-50 kg) launch demand globally in future years. Nanosatellites (NanoSats) are small spacecraft in the 1-10 kg range, which present a simple, low-cost option for developing quickly-deployable satellites. CubeSats, a special category of NanoSats, are even being considered for interplanetary missions. However, the small dimensions of CubeSats and the limited mass of the NanoSat class in general place limits of capability on their electrical power systems (especially where typical power sources such as solar panels are considered) and stored energy reserves; restricting the power budget and overall functionality. For example, leveraging NanoSat clusters for computationally intensive problems that are solved collectively becomes more challenging with power related restrictions on communication and data-processing. Further, interplanetary missions that would take NanoSats far from the sun, make the use of solar panels less effective as a power source as their required area would become quite large. To overcome these limitations, americium 241 (Am-241) has been suggested as a low power source option. The Idaho National Laboratory, Center for Space Nuclear Research reports that:  (Production) requires small quantities of isotope - 62.5 g of Pu-238; 250 g Am- 241 (for 5 We); Am-241 is available at around 1 kg/yr commercially; Am-241 produces 59 kev gammas which are stopped readily by tungsten so the radiation field is very low. Whereby, an Am-241 source could be placed in among the instruments and the waste heat used to heat the platform; and  amounts of isotope are so low that launch approval may be easier, especially with tungsten encapsulation. As further reported, Am-241 has a half-life that is approximately five times greater than that of Pu- 238 and it has been determined that the neutron yield of a 241-AmO(sub 2) source is approximately an order of magnitude lower than that of a 238-PuO(sub 2) source of equal mass and degree of (sup 16)O enrichment. Also it has been demonstrated that shielded heat sources fuelled by oxygen-enriched 238-PuO(sub 2) have masses that are up to 10 times greater than those fuelled by oxygenenriched 241-AmO(sub 2) with equivalent thermal power outputs and neutron dose rates at 1 m radii. For these reasons, Am-241 is well suited to missions that demand long duration electrical power output, such as deep spaceflight missions and similar missions that use radiation-hard electronics and instrumentation that are less susceptible to neutron radiation damage.

Robertson, Glen A.↗

NASA NTRS1960

Problems in Atmosphere Entry from Parabolic Orbits

Up to the present time we have been concerned with sending small instrumented probes into interplanetary space and larger bodies into near-earth satellite orbits. Only recently we have begun to recover intact some of these orbiting bodies and soon we will carry on this program to the recovery of manned satellites. In the not-too-distant future we will be concerned with transit to and from the moon and the planets. It is with the problems of entry into our atmosphere of these higher speed vehicles that this paper will be concerned.

Allen, H. Julian↗

NASA NTRS1969

Solar protons E greater than 100 Mev incident over Antarctica during January- February 1967.

Commencing at 0825 +a -1 UT on January 28, 1967, a large and prolonged increase in the intensity of penetrating charged particles was observed by balloon-borne instruments floating over Byrd Station, Antarctica. (80°S, 120°W). A peak intensity of approximately 50 protons per cm 2 -sec-steradian with E> 100 MeV occurred at about 1230 UT on the 28th. The event was under observation almost continuously over a period of about 100 hours until the intensity decayed below cosmic-ray background on February 1. The initial decay was rapid but, some 40 hours after onset, went over into a slow exponential decay characterized by a 20 hour time-constant. The decay phase of an additional, though considerably less intense, event was observed on February 3 and 4. Presumably both events had their origins in major disturbances on the far side of the sun since nether event has been definitely linked to any feature which existed on the visible disk within an appropriate time interval. Results pertaining to the time-intensity profile and to the energy spectrum for protons E> 100 MeV are presented for the January 29 event. Comparison of the balloon results with neutron-monitor and satellite measurements and with models of interplanetary diffusion has led to some conclusions regarding the role of small-angle scattering by irregularities and by the random walk of magnetic lines of force relative to the mean interplanetary field within the orbit of earth.

Energy spectrum↗

NASA NTRS1979

Helios 2-Vela-Ariel 5 gamma-ray burst source position

The gamma-ray burst of January 28, 1976, one of 18 events thus far detected in interplanetary space with Helios 2, was also observed with the Vela 5A and 6A and the Ariel 5 satellites. A small source field is obtained from the intersection of the region derived from the observed time delays between Helios 2 and Vela 5A and 6A, with the source region independently found with the Ariel 5 X-ray detector. This area contains neither any steady X-ray source as scanned by HEAO 1 nor any previously cataloged X-ray, radio, or infrared sources, X-ray transients, quasars, Seyferts, globular clusters, flare stars, pulsars, white dwarfs, or high enery gamma-ray sources. The region is, however, within the source field of a gamma-ray transient observed in 1974 by Jacobson et al. (1978) which exhibited nuclear gamma-ray line structure.

Cline, T. L.↗

NASA NTRS1979

Helios-2 Vela-Ariel-5 gamma-ray burst source position

The gamma-ray burst of 28 January 1976, one of 18 events thus far detected in interplanetary space with Helios-2, was also observed with the Vela-5A, -6A and the Ariel-5 satellites. A small source field is obtained from the intersection of the region derived from the observed time delays between Helios-2 and Vela-5A and -6A with the source region independently found with the Ariel-5 X-ray detector. This area contains neither any steady X-ray source as scanned by HEAO-A nor any previously catalogued X-ray, radio or infrared sources, X-ray transients, quasars, seyferts, globular clusters, flare stars, pulsars, white dwarfs or high energy gamma-ray sources. The region is however, within the source field of a gamma-ray transient observed in 1974, which exhibited nuclear gamma-ray line structure.

Cline, T. L.↗

NASA NTRS1999

Interplanetary Microlaser Transponders

The feasibility of an asynchronous (i.e. independently firing) interplanetary laser transponder, capable of ranging between Earth and Mars and using the automated SLR2000 Satellite Laser Ranging (SLR) system as an Earth base station, has been suggested. Since that time, we have received a small amount of discretionary funding to further explore the transponder concept and to develop and test an engineering breadboard. Candidate operational scenarios for acquiring and tracking the opposite laser terminal over interplanetary distances have been developed, and breadboard engineering parameters were chosen to reflect the requirements of an Earth-Mars link Laboratory tests have been devised to simulate the Earth- Mars link between two independent SLR2000 transceivers and to demonstrate the transfer of range and time in single photon mode. The present paper reviews the concept of the asynchronous microlaser transponder, the transponder breadboard design, an operational scenario recently developed for an asteroid rendezvous, and the laboratory test setup. The optical head of the transponder breadboard fits within a cylinder roughly 15 cm in diameter and 32 cm in length and is mounted in a commercial two axis gimbal driven by two computer-controlled stepper motors which allows the receiver optical axis to be centered on a simulated Earth image. The optical head is built around a small optical bench which supports a 14.7 cm diameter refractive telescope, a prototype 2 kHz SLR2000 microlaser transmitter, a quadrant microchannel plate photomultiplier (MCP/PMT), a CCD array camera, spatial and spectral filters, assorted lenses and mirrors, and protective covers and sun shields. The microlaser is end-pumped by a fiber-coupled diode laser array. An annular mirror is employed as a passive transmit/receive (T/R) switch in an aperture-sharing arrangement wherein the transmitted beam passes through the central hole and illuminates only the central 2.5 cm of the common telescope (adequate to achieve a 10 arcsecond full laser beam divergence) while the receiver uses the remainder of the 14.7 cm aperture. Additional electronic instrumentation includes the diode pump array and associated heat sink and current drivers, rubidium frequency standard, timing distribution module, range gate generator, a recently developed all-digital correlation range receiver, and system computer. Acquisition of the opposite transponder terminal requires a search within a three-dimensional volume determined by the initial pointing uncertainty and a maximum 500 microsecond uncertainty in the laser time of fire at the opposite terminal for totally uncorrelated Earth and spacecraft clocks. The angular search is aided by a sensitive CCD array capable of imaging the Earth, Moon, and surrounding stars within the nominal + 0.5 degree cone of uncertainty associated with the initial pointing of a spacecraft body or microwave communications dish.

Degnan, John J.↗

NASA NTRS2016

MEMS Reaction Control and Maneuvering for Picosat Beyond LEO

The MEMS Reaction Control and Maneuvering for Picosat Beyond LEO project will further develop a multi-functional small satellite technology for low-power attitude control, or orientation, of picosatellites beyond low Earth orbit (LEO). The Film-Evaporation MEMS Tunable Array (FEMTA) concept initially developed in 2013, is a thermal valving system which utilizes capillary forces in a microchannel to offset internal pressures in a bulk fluid. The local vapor pressure is increased by resistive film heating until it exceeds meniscus strength in a nozzle which induces vacuum boiling and provides a stagnation pressure equal to vapor pressure at that point which is used for propulsion. Interplanetary CubeSats can utilize FEMTA for high slew rate attitude corrections in addition to desaturating reaction wheels. The FEMTA in cooling mode can be used for thermal control during high-power communication events, which are likely to accompany the attitude correction. Current small satellite propulsion options are limited to orbit correction whereas picosatellites are lacking attitude control thrusters. The available attitude control systems are either quickly saturated reaction wheels or movable high drag surfaces with long response times.

Low-power attitude control↗

NASA NTRS2011

Estimation of Enceladus Plume Density Using Cassini Flight Data

Wang, Eric K.↗

NASA NTRS2010

Estimation and Modeling of Enceladus Plume Jet Density Using Reaction Wheel Control Data

The Cassini spacecraft was launched on October 15, 1997 by a Titan 4B launch vehicle. After an interplanetary cruise of almost seven years, it arrived at Saturn on June 30, 2004. In 2005, Cassini completed three flybys of Enceladus, a small, icy satellite of Saturn. Observations made during these flybys confirmed the existence of a water vapor plume in the south polar region of Enceladus. Five additional low-altitude flybys of Enceladus were successfully executed in 2008-9 to better characterize these watery plumes. The first of these flybys was the 50-km Enceladus-3 (E3) flyby executed on March 12, 2008. During the E3 flyby, the spacecraft attitude was controlled by a set of three reaction wheels. During the flyby, multiple plume jets imparted disturbance torque on the spacecraft resulting in small but visible attitude control errors. Using the known and unique transfer function between the disturbance torque and the attitude control error, the collected attitude control error telemetry could be used to estimate the disturbance torque. The effectiveness of this methodology is confirmed using the E3 telemetry data. Given good estimates of spacecraft's projected area, center of pressure location, and spacecraft velocity, the time history of the Enceladus plume density is reconstructed accordingly. The 1 sigma uncertainty of the estimated density is 7.7%. Next, we modeled the density due to each plume jet as a function of both the radial and angular distances of the spacecraft from the plume source. We also conjecture that the total plume density experienced by the spacecraft is the sum of the component plume densities. By comparing the time history of the reconstructed E3 plume density with that predicted by the plume model, values of the plume model parameters are determined. Results obtained are compared with those determined by other Cassini science instruments.

Lee, Allan Y.↗

NASA NTRS2010

Estimation and Modeling of Enceladus Plume Jet Density Using Reaction Wheel Control Data

The Cassini spacecraft was launched on October 15, 1997 by a Titan 4B launch vehicle. After an interplanetary cruise of almost seven years, it arrived at Saturn on June 30, 2004. In 2005, Cassini completed three flybys of Enceladus, a small, icy satellite of Saturn. Observations made during these flybys confirmed the existence of a water vapor plume in the south polar region of Enceladus. Five additional low-altitude flybys of Enceladus were successfully executed in 2008-9 to better characterize these watery plumes. The first of these flybys was the 50-km Enceladus-3 (E3) flyby executed on March 12, 2008. During the E3 flyby, the spacecraft attitude was controlled by a set of three reaction wheels. During the flyby, multiple plume jets imparted disturbance torque on the spacecraft resulting in small but visible attitude control errors. Using the known and unique transfer function between the disturbance torque and the attitude control error, the collected attitude control error telemetry could be used to estimate the disturbance torque. The effectiveness of this methodology is confirmed using the E3 telemetry data. Given good estimates of spacecraft's projected area, center of pressure location, and spacecraft velocity, the time history of the Enceladus plume density is reconstructed accordingly. The 1-sigma uncertainty of the estimated density is 7.7%. Next, we modeled the density due to each plume jet as a function of both the radial and angular distances of the spacecraft from the plume source. We also conjecture that the total plume density experienced by the spacecraft is the sum of the component plume densities. By comparing the time history of the reconstructed E3 plume density with that predicted by the plume model, values of the plume model parameters are determined. Results obtained are compared with those determined by other Cassini science instruments.

Lee, Allan Y.↗

NASA NTRS1985

The distant magnetotail

Satellite, balloon, aircraft, and ground-based data are presented for the night of February 24, 1974, when variations in the interplanetary magnetic field caused a series of small substorms. The data are interpreted as indicating that a series of small plasmoids were formed in the magnetotail, travelled tailward rubbing on closed field lines, and produced poleward travelling auroral features as the ionospheric footprint of the plasmoids. When the plasmoids cross the X-line and magnetopause onto interplanetary magnetic field lines they lose connection with the earth and the auroral features disappear. In this instance, the closed field line portion reached up to nearly 80 deg in geomagnetic latitude, over 10 deg poleward from the auroral oval.

Heikkila, W.↗

NASA NTRS2003

Structure of the Outer Cusp and Sources of the Cusp Precipitation during Intervals of a Horizontal IMF

The cusp represents a place where the magnetosheath plasma can directly penetrate into the magnetosphere. Since the main transport processes are connected with merging of the interplanetary and magnetospheric field lines, the interplanetary magnetic field (IMF) Orientation plays a decisive role in the formation of the high-altitude cusp. The importance of the sign of the IMF Bz component for this process was suggested about 40 years ago and later it was documented by many experimental investigations. However, situations when IMF Bz is the major IMF component are rather rare. The structure of the cusp during periods of a small IMF BZ is generally unknown, probably due to the fully 3-D nature of the interaction. The present case study reveals the importance of horizontal IMF components on the global magnetospheric configuration as well as on small-scale processes at the cusp-magnetosheath interface. We have used simultaneous measurements of several spacecraft (ISTP program) operating in different regions of interplanetary space and two closely spaced satellites (INTERBALL-1/MAGION- 4) crossing the cusp-magnetosheath boundary to show the connection between the short- and large-scale phenomena. In the northern hemisphere, observations suggest a presence of two spots of cusp-like precipitation supplied by reconnection occurring simultaneously in both hemispheres. A source of this bifurcation is the positive IMF By component further enhanced by the field draping in the magnetosheath. This magnetic field component shifts the entry point far away from the local noon but in opposite sense in either hemisphere.

Nemecek, Z.↗

NASA NTRS2003

Structure of the Outer Cusp and Sources of the Cusp Precipitation during Intervals of a Horizontal IMF

The cusp represents a place where the magnetosheath plasma can directly penetrate into the magnetosphere. Since the main transport processes are connected with merging of the interplanetary and magnetospheric field lines: the interplanetary magnetic field (IMF) Orientation plays a decisive role in the formation of the high-altitude cusp. The importance of the sign of the IMF B(sub Z) component for this process was suggested about 40 years ago and later it was documented by many experimental investigations. However, situations when IMF Bz is the major IMF component are rather rare. The structure of the cusp during periods of a small IMF B(sub Z) is generally unknown, probably due to the fully 3-D nature of the interaction. The present case study reveals the importance of horizontal IMF components on the global magnetospheric configuration as well as on small-scale processes at the cusp-magnetosheath interface. We have used simultaneous measurements of several spacecraft (ISTP program) operating in different regions of interplanetary space and two closely spaced satellites (INTERBALL-1/MAGION-4) crossing the cusp-magnetosheath boundary to show the connection between the short- and large-scale phenomena. In the northern hemisphere, observations suggest a presence of two spots of cusp-like precipitation supplied by reconnection occurring simultaneously in both hemispheres. A source of this bifurcation is the positive IMF B(sub y) component further enhanced by the field draping in the magnetosheath. This magnetic field component shifts the entry point far away from the local noon but in opposite sense in either hemisphere. The cusp represents a place where the magnetosheath plasma can directly

Berchem, Jean↗

NASA NTRS2000

Svalbard: A Window for Understanding Temporal/Spatial Aspects of Solar Wind Coupling to the Magnetosphere and Ionosphere

The research facilities on the Svalbard archepelego provide a unique opportunity for observing the temporal and spatial characteristics of the interaction of the solar wind with the magnetosphere-ionosphere system. The first sounding rockets from the SvdRak range at Ny-Alesund have opened a new perspective for separating temporal and spatial effects. Finding a correlation between observations from the rockets and the Wind satellite located nearly 200 R(sub E) upstream in the solar wind that had a lag time less than the advection time forced consideration of tilted phase planes of the interp1anetary electric field. From this it was deduced that the interaction process has to involve high latitude merging, the interplanetary magnetic field (IMF) B(sub Y) bifurcates the cusp relative to the high latitude source regions in the Northern and Southern Hemispheres, IMF B(sub X) controls the interaction time in each hemisphere, and the small convection cell is driven by opposite hemisphere merging. Studies with 4 satellites in the solar wind have shown that the tilt of the phase plane varies on a minute-by-minute basis and change significantly on time scales of tens of minutes. Images of the cusp at 557.7 nm provide a nearly instantaneous picture of the temporal and spatial characteristics of merging at the magnetopause and show that the process can occur at multiple locations. The rate and location vary with time.

Maynard, Nelson C.↗