Session VII: Advanced Technologies 1

Chair: Rory Barrett – Composite Technology Development, Inc.

Wednesday, August 13, 2008

8:30 a.m.Integrated Optically Transparent Solar Cell Antennas Made from Meshed Conductors

Timothy Turpin, Reyhan Baktur – Utah State University

ABSTRACT: This paper presents the feasibility of integrating meshed patch antennas directly onto the solar cell assembly to save valuable surface real estate of a small satellite. The solar cell cover glass is used as the substrate for the patch antenna. The integrated patch antennas are required to be transparent to light in order to ensure the proper operation of solar cells.

It is found that the mesh geometry can be designed to offer the optimal transparency and antenna radiation at the same time. To design the transparent solar cell antennas in the presence of photovoltaic cells, Ansoft’s HFSS is used to model patch antennas on solar cells and the dielectric constant of solar cells is approximated with that of silicon.

In order to verify the design, meshed antennas are printed with conductive ink on plastic substrate and measured results are compared against the design data from simulations.

8:45 a.m. On–Orbit Spacecraft Inertia and Rate Sensor Scale Factor Estimation for Microsatellites: Philip Ferguson – Microsat Systems Canada Incorporated (MSCI); ABSTRACT: A key challenge in testing and operating small satellites is the determination of the moment of inertia. Attitude control systems engineers use the moment of inertia to develop closed loop pointing controllers as well as accurate feed-forward pointing commands that predict the satellite’s motion. Traditionally, engineers measure the satellite moment of inertia using a mass properties table. However, for small, relatively lightweight satellites, this process is error-prone and costly regardless of the satellite size. This paper presents a novel on-orbit inertia-estimation technique. The algorithm is based on standard non-linear function solvers that can be run on the ground and requires only a rudimentary initial inertia estimate as a starting point (such an estimate can be obtained from structural modeling software). In addition to estimating the satellite inertia matrix, the estimator can also provide rate sensor scale-factor corrections. This paper demonstrates the inertia and scale factor estimator using the MOST spacecraft (now in its fifth year of operations).

9:00 a.m.A Novel Cold Gas Propulsion System for Nanosatellites and Picosatellites: David Hinkley – The Aerospace Corporation; ABSTRACT: A Microelectromechanical System-based (MEMS) PICOSAT Inspector (MEPSI) picosatellite on STS-116, in December 2006, used a five thruster cold gas propulsion system to translate and rotate. The inspector picosatellite measured 4 x 4 x 5 inches in dimension and weighted 1.4 kg. Our propulsion system was produced by a unique (to spacecraft) method of manufacturing that is low cost, tightly integrated, and leak tight. This paper will describe the design, fabrication, testing and limits of this type of unit, and extrapolate to other related uses found at The Aerospace Corporation. This work was funded by The Aerospace Corporations Independent Research and Development (IR&D) program.

9:15 a.m.A Compact Low–Power High–Isp Thruster for Microsatellites: Douglas Spence, Nathaniel Demmons, Thomas Roy – Busek Co., Inc.; ABSTRACT: Busek Co. Inc. completed delivery of flight-qualified colloid thrusters for NASAs ST7 mission in May 2008. This effort has led to development of variants of the technology suitable for small satellite applications. Colloid thrusters operate by electrostatically accelerating charged droplets of an electrically conductive ionic liquid, and are capable of providing a high degree of throttling and variable Isp. Life tests of the ST7 thrusters have demonstrated over 3000 hours of continuous operation with no deterioration in performance. A further benefit is that the colloid thrusters do not present high pressure and fire safety hazards common to many other propulsion systems- the propellant is nonreactive and is typically stored at less than 20psig. The thrusters presented have a target maximum thrust of 1 milliNewton with 0.1-1.0 milliNewton throttling. They are designed to operate in the Isp range of 400-1000s, consume a maximum of 15W (including power supply losses), and be self-contained in a 10cm x 10cm x 20cm package requiring only power and thrust command inputs. The package contains sufficient propellant for 500 hours operation at maximum thrust, yielding total impulse of 1800 seconds capable of imparting almost 200 m/s delta V to a 10kg satellite.

9:30 a.m.Development of a Micro–Thruster Impulse Measurement System Using Optical Sensors: Young–Keun Chang, –Jin Kang, Hyea–Ran Cho – Korea Aerospace University; ABSTRACT: A new method for measuring performance of a micro-thruster is suggested in this paper. A few thrust stands have been developed for measuring micro-level thrusts. This paper describes a different measurement method that can minimize the calibration involved in the measurements, while providing the capability of directly measuring the produced minimum impulse bit. The underlying theory and the theoretical background for the measurement mechanism are described here. The theory and method is verified using computer simulation, and the result is given in this paper. The theory has also been tested on an actual hardware. The prototype measurement system has been tested inside a vacuum chamber for verification of the theoretical and simulation results. Actual experimental data was used to verify the theory, and a test cold gas thruster was also employed for final testing and verification of the measurement system.

9:45 a.m.Micro/Nanotechnology for Picosatellites: Siegfried Janson – The Aerospace Corporation; ABSTRACT: Up until the year 2000, only a few active picosatellites had been put into orbit. For the first 40 years of the Space Age, it was difficult to integrate high levels of functionality into the picosatellite 0.1 to 1-kg mass range. Fortunately, continuing advancements in micro/nanoelectronics and microelectromechanical systems has now enabled many nanosatellite and microsatellite capabilities to be implemented in picosatellites. Complementary metal oxide semiconductor (CMOS) micro/nanoelectronics are currently mass-produced with lateral structures smaller than 65-nm, thus enabling creation of billion-transistor integrated circuits on cm-scale silicon dice. Microelectromechanical systems (MEMS) are fabricated using similar processes and will benefit from further reductions in minimum feature size over time. Micro/nanoelectronics and micro/nanoelectromechanical systems will evolve over the next decade to provide ever-higher levels of functional density per unit area. Small spacecraft, particularly picosatellites and CubeSats, require mm-to-cm scale sensors for attitude determination. Commercial CMOS technology provides mm-to-cm scale image sensors that can function as sun and star sensors while MEMS technology offers mm-to-cm scale magnetic and inertial sensors. Custom CMOS/MEMS technology enables mm-scale sun and horizon sensors suitable for picosatellites and even smaller spacecraft. Several examples of millimeter and centimeter-scale sun sensors are given.

AlternateSatellite Modular and Reconfigurable Thermal System (SMARTS): David Bugby – ATK Space; Walter Zimbeck – Technology Assessment & Transfer, Inc.; Edward Kroliczek – B&K Engineering; Andrew Williams – Air Force Research Laboratory/Space Vehicles Directorate; ABSTRACT: This paper describes a two-phase heat transfer based thermal design architecture for satellites that need to be conceived, configured, launched, and operationally deployed very quickly. The architecture has been given the acronym SMARTS for Satellite Modular and Reconfigurable Thermal System. SMARTS is a Phase I-II SBIR program awarded by the Air Force Research Laboratory, Space Vehicles Directorate to Technology Assessment & Transfer. The SMARTS philosophy involves four basic design rules: (1) modest radiator oversizing; (2) maximum external insulation; (3) internal isothermalization; and (4) radiator heat-flow modulation. For a prototypical multi-panel small satellite, the paper describes a SMARTS thermal control system that uses: (a) panel-to-panel heat conduction; (b) intra-panel heat pipe isothermalization; (c) radiator heat-flow modulation via a thermoelectric cooler (TEC) cold-biased loop heat pipe (LHP); and (d) maximum external MLI. Analyses are presented that compare the traditional "cold-biasing plus heater power" passive thermal design approach to the SMARTS approach. Additional analyses and conceptual design work oriented towards the Phase II goal of developing a multi-panel, TEC-cold-biased, LHP-modulated SMARTS small satellite test bed are also described. The ultimate goal is to incorporate SMARTS into the design of future satellites envisioned by the Operationally Responsive Space (ORS) initiative.

AlternateImplications of Advanced Thermal Control Architecture for Modular Spacecraft: Quinn Young – Space Dynamics Laboratory; Andrew Williams – Air Force Research Laboratory; Brent Stucker – Utah State University; ABSTRACT: ABSTRACT: The combined knowledge-bases of architecture theory, thermal design, and advanced technologies have been used to develop a modular approach to thermal control. This new approach was designed specifically for the requirements of a modular spacecraft, reducing or eliminating the system-level coupling and interdependencies of traditional thermal designs. The decoupling of the thermal subsystem allows improved control methods using advanced technologies that control the heat leaving the spacecraft. These control methods eliminate the need for survival heaters, reducing mass and power as well as simplifying the analysis effort required for verification and validation of the design. In addition, the modular approach allows fundamentally different designs to be incorporated into the spacecraft architecture and structure. The modular thermal approach, structural implications, and system implications are explained with illustrations and examples.


		22nd Annual Conference on Small Satellites Technical Sessions Session VII: Advanced Technologies 1 Chair: Rory Barrett – Composite Technology Development, Inc. Wednesday, August 13, 2008 8:30 a.m.Integrated Optically Transparent Solar Cell Antennas Made from Meshed Conductors Timothy Turpin, Reyhan Baktur – Utah State University ABSTRACT: This paper presents the feasibility of integrating meshed patch antennas directly onto the solar cell assembly to save valuable surface real estate of a small satellite. The solar cell cover glass is used as the substrate for the patch antenna. The integrated patch antennas are required to be transparent to light in order to ensure the proper operation of solar cells. It is found that the mesh geometry can be designed to offer the optimal transparency and antenna radiation at the same time. To design the transparent solar cell antennas in the presence of photovoltaic cells, Ansoft’s HFSS is used to model patch antennas on solar cells and the dielectric constant of solar cells is approximated with that of silicon. In order to verify the design, meshed antennas are printed with conductive ink on plastic substrate and measured results are compared against the design data from simulations. 8:45 a.m. On–Orbit Spacecraft Inertia and Rate Sensor Scale Factor Estimation for Microsatellites Philip Ferguson – Microsat Systems Canada Incorporated (MSCI) ABSTRACT: A key challenge in testing and operating small satellites is the determination of the moment of inertia. Attitude control systems engineers use the moment of inertia to develop closed loop pointing controllers as well as accurate feed-forward pointing commands that predict the satellite’s motion. Traditionally, engineers measure the satellite moment of inertia using a mass properties table. However, for small, relatively lightweight satellites, this process is error-prone and costly regardless of the satellite size. This paper presents a novel on-orbit inertia-estimation technique. The algorithm is based on standard non-linear function solvers that can be run on the ground and requires only a rudimentary initial inertia estimate as a starting point (such an estimate can be obtained from structural modeling software). In addition to estimating the satellite inertia matrix, the estimator can also provide rate sensor scale-factor corrections. This paper demonstrates the inertia and scale factor estimator using the MOST spacecraft (now in its fifth year of operations). 9:00 a.m.A Novel Cold Gas Propulsion System for Nanosatellites and Picosatellites David Hinkley – The Aerospace Corporation ABSTRACT: A Microelectromechanical System-based (MEMS) PICOSAT Inspector (MEPSI) picosatellite on STS-116, in December 2006, used a five thruster cold gas propulsion system to translate and rotate. The inspector picosatellite measured 4 x 4 x 5 inches in dimension and weighted 1.4 kg. Our propulsion system was produced by a unique (to spacecraft) method of manufacturing that is low cost, tightly integrated, and leak tight. This paper will describe the design, fabrication, testing and limits of this type of unit, and extrapolate to other related uses found at The Aerospace Corporation. This work was funded by The Aerospace Corporations Independent Research and Development (IR&D) program. 9:15 a.m.A Compact Low–Power High–Isp Thruster for Microsatellites Douglas Spence, Nathaniel Demmons, Thomas Roy – Busek Co., Inc. ABSTRACT: Busek Co. Inc. completed delivery of flight-qualified colloid thrusters for NASAs ST7 mission in May 2008. This effort has led to development of variants of the technology suitable for small satellite applications. Colloid thrusters operate by electrostatically accelerating charged droplets of an electrically conductive ionic liquid, and are capable of providing a high degree of throttling and variable Isp. Life tests of the ST7 thrusters have demonstrated over 3000 hours of continuous operation with no deterioration in performance. A further benefit is that the colloid thrusters do not present high pressure and fire safety hazards common to many other propulsion systems- the propellant is nonreactive and is typically stored at less than 20psig. The thrusters presented have a target maximum thrust of 1 milliNewton with 0.1-1.0 milliNewton throttling. They are designed to operate in the Isp range of 400-1000s, consume a maximum of 15W (including power supply losses), and be self-contained in a 10cm x 10cm x 20cm package requiring only power and thrust command inputs. The package contains sufficient propellant for 500 hours operation at maximum thrust, yielding total impulse of 1800 seconds capable of imparting almost 200 m/s delta V to a 10kg satellite. 9:30 a.m.Development of a Micro–Thruster Impulse Measurement System Using Optical Sensors Young–Keun Chang, –Jin Kang, Hyea–Ran Cho – Korea Aerospace University ABSTRACT: A new method for measuring performance of a micro-thruster is suggested in this paper. A few thrust stands have been developed for measuring micro-level thrusts. This paper describes a different measurement method that can minimize the calibration involved in the measurements, while providing the capability of directly measuring the produced minimum impulse bit. The underlying theory and the theoretical background for the measurement mechanism are described here. The theory and method is verified using computer simulation, and the result is given in this paper. The theory has also been tested on an actual hardware. The prototype measurement system has been tested inside a vacuum chamber for verification of the theoretical and simulation results. Actual experimental data was used to verify the theory, and a test cold gas thruster was also employed for final testing and verification of the measurement system. 9:45 a.m.Micro/Nanotechnology for Picosatellites Siegfried Janson – The Aerospace Corporation ABSTRACT: Up until the year 2000, only a few active picosatellites had been put into orbit. For the first 40 years of the Space Age, it was difficult to integrate high levels of functionality into the picosatellite 0.1 to 1-kg mass range. Fortunately, continuing advancements in micro/nanoelectronics and microelectromechanical systems has now enabled many nanosatellite and microsatellite capabilities to be implemented in picosatellites. Complementary metal oxide semiconductor (CMOS) micro/nanoelectronics are currently mass-produced with lateral structures smaller than 65-nm, thus enabling creation of billion-transistor integrated circuits on cm-scale silicon dice. Microelectromechanical systems (MEMS) are fabricated using similar processes and will benefit from further reductions in minimum feature size over time. Micro/nanoelectronics and micro/nanoelectromechanical systems will evolve over the next decade to provide ever-higher levels of functional density per unit area. Small spacecraft, particularly picosatellites and CubeSats, require mm-to-cm scale sensors for attitude determination. Commercial CMOS technology provides mm-to-cm scale image sensors that can function as sun and star sensors while MEMS technology offers mm-to-cm scale magnetic and inertial sensors. Custom CMOS/MEMS technology enables mm-scale sun and horizon sensors suitable for picosatellites and even smaller spacecraft. Several examples of millimeter and centimeter-scale sun sensors are given. AlternateSatellite Modular and Reconfigurable Thermal System (SMARTS) David Bugby – ATK Space; Walter Zimbeck – Technology Assessment & Transfer, Inc.; Edward Kroliczek – B&K Engineering; Andrew Williams – Air Force Research Laboratory/Space Vehicles Directorate ABSTRACT: This paper describes a two-phase heat transfer based thermal design architecture for satellites that need to be conceived, configured, launched, and operationally deployed very quickly. The architecture has been given the acronym SMARTS for Satellite Modular and Reconfigurable Thermal System. SMARTS is a Phase I-II SBIR program awarded by the Air Force Research Laboratory, Space Vehicles Directorate to Technology Assessment & Transfer. The SMARTS philosophy involves four basic design rules: (1) modest radiator oversizing; (2) maximum external insulation; (3) internal isothermalization; and (4) radiator heat-flow modulation. For a prototypical multi-panel small satellite, the paper describes a SMARTS thermal control system that uses: (a) panel-to-panel heat conduction; (b) intra-panel heat pipe isothermalization; (c) radiator heat-flow modulation via a thermoelectric cooler (TEC) cold-biased loop heat pipe (LHP); and (d) maximum external MLI. Analyses are presented that compare the traditional "cold-biasing plus heater power" passive thermal design approach to the SMARTS approach. Additional analyses and conceptual design work oriented towards the Phase II goal of developing a multi-panel, TEC-cold-biased, LHP-modulated SMARTS small satellite test bed are also described. The ultimate goal is to incorporate SMARTS into the design of future satellites envisioned by the Operationally Responsive Space (ORS) initiative. AlternateImplications of Advanced Thermal Control Architecture for Modular Spacecraft Quinn Young – Space Dynamics Laboratory; Andrew Williams – Air Force Research Laboratory; Brent Stucker – Utah State University ABSTRACT: ABSTRACT: The combined knowledge-bases of architecture theory, thermal design, and advanced technologies have been used to develop a modular approach to thermal control. This new approach was designed specifically for the requirements of a modular spacecraft, reducing or eliminating the system-level coupling and interdependencies of traditional thermal designs. The decoupling of the thermal subsystem allows improved control methods using advanced technologies that control the heat leaving the spacecraft. These control methods eliminate the need for survival heaters, reducing mass and power as well as simplifying the analysis effort required for verification and validation of the design. In addition, the modular approach allows fundamentally different designs to be incorporated into the spacecraft architecture and structure. The modular thermal approach, structural implications, and system implications are explained with illustrations and examples.