A space-to-space microwave wireless power transmission experiential mission using small satellites

Acta Astronautica 103 (2014) 193–203

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A space-to-space microwave wireless power transmission experiential mission using small satellites Corey Bergsrud a,n, Jeremy Straub b,1 a Department of Electrical Engineering, University of North Dakota, Upson Hall, IIRoom 160 243 Centennial Drive, Stop 7165 Grand Forks, ND 58202-7165, USA b Department of Computer Science, University of North Dakota, Streibel Hall, Room 201 3950 Campus Road, Stop 9015, Grand Forks, ND 58202-9015, USA

a r t i c l e i n f o

abstract

Article history: Received 20 March 2014 Received in revised form 13 May 2014 Accepted 23 June 2014 Available online 2 July 2014

A space solar microwave power transfer system (SSMPTS) may represent a paradigm shift to how space missions in Earth orbit are designed. A SSMPTS may allow a smaller receiving surface to be utilized on the receiving craft due to the higher-density power transfer (compared to direct solar flux) from a SSMPTS supplier craft; the receiving system is also more efficient and requires less mass and volume. The SSMPTS approach also increases mission lifetime, as antenna systems do not degrade nearly as quickly as solar panels. The SSMPTS supplier craft (instead) can be replaced as its solar panels degrade, a mechanism for replacing panels can be utilized or the SSMPTS can be maneuvered closer to a subset of consumer spacecraft. SSMPTS can also be utilized to supply power to spacecraft in eclipse and to supply variable amounts of power, based on current mission needs, to power the craft or augment other power systems. A minimal level of orbital demonstrations of SSP technologies have occurred. A mission is planned to demonstrate and characterize the efficacy of space-to-space microwave wireless power transfer. This paper presents an overview of this prospective mission. It then discusses the spacecraft system (comprised of an ESPA/SmallSat-class spacecraft and a 1-U CubeSat), launch options, mission operations and the process of evaluating mission outcomes. Published by Elsevier Ltd. on behalf of IAA.

Keywords: Space solar power satellite Space-to-space power transfer Microwave power transmission Wireless power transmission Experimental mission design

1. Introduction Conventional space missions operate either from an onboard nuclear power source or using onboard solar power generation panels. In the former case, the mission launches with all of the power that it will ever have; in the latter, power utilization is (similarly) limited by generation

n Corresponding author. Tel.: þ1 (701) 777 4331; fax: þ1 (701) 777 5253. E-mail addresses: [email protected] (C. Bergsrud), [email protected] (J. Straub). 1 Tel.: þ1 (701) 777 4107; fax: þ 1 (701) 777 3330.

http://dx.doi.org/10.1016/j.actaastro.2014.06.033 0094-5765/Published by Elsevier Ltd. on behalf of IAA.

capability. In both cases, the amount of power being produced declines as the mission progresses. These spacecraft must, thus, be built to satisfy end-of-mission generation needs, possibly meaning that excessive power is generated early in the mission which must be radiated as heat (necessitating a more capable thermal system). The additional generation capability (as well as, potentially, the expanded thermal system) increases spacecraft volume, mass, complexity and costs. The deterioration effectively limits the spacecraft's lifespan. A new approach has been proposed [1–4] which utilizes space solar power (SSP) satellites to provide power to other spacecraft under a utility-provider model, using

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small spacecraft. Utilizing this approach and an orbital service model [5,6], consumers can procure power for their spacecraft on an ongoing, augmentation contract or ad hoc basis. This power is then transmitted to rectifying antennas (rectennas) on their spacecraft for use. The use of rectennas instead of solar panels increase spacecraft lifespan and can serve to reduce mass and volume due to the greater energy density that can be transmitted in this way the rectenna is also lighter, for a given amount of surface area, than a solar array. Before a system of this type can be developed commercially, for government use or otherwise, the various technologies and their integrated operations must be tested and space qualified. This paper provides an overview of a mission which is designed to demonstrate the space-to-space microwave wireless power transfer concept, test key assumptions and components and begin the process of space qualifying the individual components and system operations. In the following section, background material is presented. After this, there is an overview of the mission. Next, an overview of the spacecraft system is provided. Then launch requirements and options are discussed, followed by an overview of mission operations. Finally, before concluding, a set of metrics for mission evaluation are presented. 2. Background Three areas of prior work inform the mission presented herein. First, prior work related to space solar power is discussed. Next, prospective uses of space solar power are discussed. Finally, an overview of small satellites (both at the ESPA/SmallSat and CubeSat sizes used in this mission) is presented. 2.1. Space solar power satellite systems A history of SSP is presented by Strassner and Chang [7], McSpadden and Mankins [8] and Bergsrud and Noghanian [9]. The concept of Space Solar Power Satellite (SSPS) systems was conceived by Dr. Peter Glaser at the Arthur D. Little company in 1973 [10]. Glaser's vision was to place large satellites in geostationary Earth orbit (GEO) whose sole purpose is to harvest large amounts of solar energy, transform the solar energy in microwave energy and transmit it to a rectifying antenna (rectenna) array located on Earth's surface. The rectenna array would collect the microwave energy and convert it into usable direct current power that is injected into the terrestrial electric grid system to supply humanity with a clean source of baseload electrical power. SSPS systems may one day supply sufficient amounts of electrical power to Earth and beyond to aid humanity in its continued advancement. The first major study of SSPS systems was conducted between 1976 and 1980 through a joint collaboration between the Department of Energy (DOE) and National Aeronautics and Space Administration (NASA) [11] as well as many other entities. This period of study resulted in the creation of the foundational architecture for SSPS systems. Large-scale power infrastructure in space consisting of about 60 SPSS, each delivering 5 GW of base load power

to the U.S. electrical grid [8] was crafted. In addition, candidate locations for SSPS rectifying antennas on Earth were investigated [12,13] along with rectenna-related atmospheric effects [14] and ways to effect electrostatic protection of the SSPS [15] and lightning protection for the rectenna [16]. This extremely productive period of the Satellite Power System Concept Development and Evaluation Program [8] determined that SSPS systems was a feasible technology and should be pursued in the future [7], and the U.S. National Research Council (NRC) recommended that the concept be re-assessed in about ten years, subsequent to additional technology development and maturation [8]. During the 1980s and early 1990s, international interest in the SSPS concept emerged in Japan, Europe, and Canada. In particular, a Japanese research group from Kyoto University conducted the first successful Microwave Wireless Power (MWP) transmission experiment (in 1983) called the Microwave Ionosphere Nonlinear Interaction eXperiment (MINIX) [17]. Again, in 1993, another Japanese research group completed the International Space Year-Microwave Energy Transmission in Space (ISY-METS) S-520-16 sounding rocket experiment [18]. Both projects utilized a daughter-mother rocket combination to demonstrate MWP transmission technologies and produced results characterizing the nonlinear plasma effects of the high power microwave energy beam in the space environment. NASA recognized the accomplishments of the Japanese teams and, in 1995 undertook a new study of the challenges of large-scale SSPS systems through the Fresh Look Study [19,20]. The study highlighted recent technological advancements which made SSPS systems more viable than they were at the end of the 1980s [21,22]. In 1998, NASA conducted the SSP Concept Definition Study, in which experts outside the agency were also involved. The SSP Concept Definition Study validated findings in the Fresh Look Study, but it also invalidated some earlier ideas which narrowed the SSP concepts. A key outcome of this study was the definition of a family of strategic Research and Technology (R&T) road maps for the possible development of SSP technologies [8]. The next major advancement of SSPS occurred in 2000 when the NASA Marshall Space Flight Center (MSFC) conducted the SSP Scientific Exploratory Research and Technology (SERT) program. The SERT program broadened the scientific community's involvement and resulted in successful demonstrations of a variety of system level components [7]. It included tightly focused exploratory research targets and rapid analysis to identify promising system concepts and establish their technical viability. Small scale demonstrations of key SSPS concepts/components using nearer-term technologies were initiated [8]. Finally, the SERT program addressed issues related to economic and societal assessment, environmental effects, resource requirements, and legal issues [7]. From the end of the SERT program to today, numerous articles have been written about SSPS systems, but supportive research has been sporadic (at best) due to funding limitations [7]. One MWP transmission project of particular interest was performed in Hawaii. The Hawaii MWP transmission experiment was carried out in 2008 by

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Mankins of Managed Energy Technologies in collaboration with researchers at Texas A&M University and the University of Kobe. The experiment transmitted power from Maui to Hawaii's big island, a distance of over 148 km [23]. With a new distance record reached, the amount of power received was less than 0.00001 W of the power transmitted [24]. Strassner [7] points out that from a MWP transmission standpoint this is considered a failure by many. However, it goes on to clarify that the poor performance was predicable since the transmitting and receiving arrays were far too small to allow efficient transfer over the 148 km, due to massive spillover loss levels. Furthermore, this experiment was conducted on a $1 million budget, limiting to what could be accomplished. 2.2. Space to space Uses of space solar power At least as early as 1985, researchers were considering space-to-space power transfer missions [25]. In 1996 mission concept the space transportation system (shuttle) would be utilized to beam power from an onboard phased array to a rectenna receiver on a satellite [26]. This mission sought to characterize system performance in much the way as the mission described herein seeks to. The use of space-to-space testing to facilitate space-to-Earth applications has also been proposed. The utility of space-to-space microwave wireless power transfer (S2S-MWPT) goes far beyond testing and demonstration, however. Previous work has considered the utility of utilizing space-to-space power transfer for supporting other spacecraft [27,28].Under this model [2], spacecraft operators would contract with a space power utility provider who would provide power on a constant, recurring, pre-contracted on-demand or ad-hoc basis. This power could serve as the primary power system of the spacecraft (onboard solar generation capabilities and storage could serve as a backup for mission-critical craft), as a constant augmentation power supply, or it could be utilized to facilitate having sufficient power for certain operations (e.g., orbit raising using and electronic propulsion system such as Keidar et al.[29] etc.). Details for this type of system were presented in [1,27] and its control was discussed in [30,31]. S2S-MWPT may potentially be useful for deep-space applications as well. 2.3. Small spacecraft Small spacecraft are as old as space travel itself. The first spacecraft, Sputnik, was also the first small spacecraft; the United States' first spacecraft, Explorer 1, was also a small spacecraft (by a mass definition). Both had a mass of less than 100 kg [32,33]. Swartwout [34], however, argues that it is not the size, but instead the use that should classify spacecraft. To-date, over 196 missions that fall into what he terms the “university class” have been launched from at least 98 different institutions [35]. University class missions are characterized by the student learning benefits they provide and the risk tolerance, which allows them to try non-traditional solutions to problems [34]. One class of small satellite, the CubeSat, has received particular attention. CubeSats are designed to be easy to

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construct due to their well-defined specifications [36] and small size and mass. Initially designed for student projects [37], CubeSats are now used for commercial [38], military [39,40], bona fide scientific experimentation [41–43] and other uses. The CubeSat concept was initially invented by Twiggs at Stanford [37] with Puig Suari at Cal Poly creating the P-POD launcher [44] to facilitate their easy launch vehicle integration and wide use. They have been used extensively in Earth orbit [35] and proposed for use for lunar [45] and deep space applications [46]. CubeSat and SmallSat/ESPA-class growth has been facilitated by the availability of free-to-qualified developer [47–49] launches; low-cost commercial launches [50] appear to also be on the horizon. Most small spacecraft are launched as secondary payloads, making their launch date, target orbit and other characteristics subject to the needs of the primary payload. 3. Mission overview The SSP test and demonstration mission seeks to demonstrate the efficacy of SSPsystems as well as to demonstrate and characterize S2S-MWPT. It also seeks to provide some ancillary benefits for project participants and the University of North Dakota related to university perception and educational outcomes as small spacecraft development programs have been shown, previously [51–54], to be quite effective at generating these types of learning outcomes. The primary, mission objectives is to demonstrate the viability of in space microwave power transfer and collect data that enables future innovation in this area. As a multi-spacecraft mission, the mission architecture can be sub-divided into two sub-missions: one for the transmitting spacecraft and one for the receiving spacecraft, which is ejected from the larger transmitting spacecraft. The mission architecture details for the transmitting spacecraft are presented in Table 1. The details for the receiving spacecraft are presented in Table 2. The transmitting spacecraft will be the mission's primary spacecraft and the sub-spacecraft (the 1-U receiving CubeSat) will be ejected from this craft. The two spacecraft will be launched as a secondary payload into an orbit with an altitude between 250 km and 800 km and an inclination of at least 47.911 (so as to overfly the University of North Dakota's ground station). Several possibilities for suitable orbits exist, including a sun-synchronous orbit (which would allow constant generation to maximize the power available for experimentation). The primary spacecraft will fly as a secondary payload on a launch with a larger satellite as its primary payload or, potentially, it could fly on a peer-primary mission on a smaller class launch vehicle. Shortly after being ejected from the launch vehicle, it will deploy its solar panels and antenna system. It will also eject the receiving subsatellite, a 1-U CubeSat, which will be allowed to move away from the primary spacecraft. This 1-U CubeSat will then deploy its rectenna array structure. The 1-U CubeSat will continue to move away for a period of time before its thrusters (powered via the transmitted power from the primary craft which is collected using the rectenna array

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Table 1 Transmitting spacecraft mission architecture details. Element

Description

Explanation

Subject

Onboard experiment

Payload

Experiment hardware and data gathering hardware

Spacecraft bus

Custom

This sub-mission will assess the capabilities and efficacy of the SSP transmitting spacecraft as well as the efficacy of the entire transmitting/system The payload is comprised of all elements required to perform the transmission and test of transmission-receiving system missions. This includes: solar power collection (solar cell) hardware, transmission hardware, a GPS and an IMU A bus will be created to facilitate efficient power collection and transfer. Key design features will include: (A) maximization of generation capability, (B) maximization of pointing, (C) reuse of structural elements for fuel storage

Launch system Orbit

To be determined by launch provider Inclination 447.911 and altitude The selected inclination facilitates overflight of Grand Forks, ND to allow communication between 250 km and 800 km with our ground station hardware. Altitude: the selection is based on communications (Earth-to-spacecraft) and orbital lifetime considerations. UND-developed A SDR-based ground station is under development at UND Direct communications between The communications will operate on the amateur frequency ranges, using an UND/satellite experimental-class license.

Ground system Communications architecture

Table 2 Receiving Spacecraft Mission Architecture Details. Element

Description

Explanation

Subject

Onboard experiment

Payload

Experiment hardware and data gathering hardware Custom

A rectenna and associated power-receiver hardware will be included. An electric thruster will be included to allow the demonstration of this operating from transmitted power. Experiment hardware includes the power receiving hardware and the thruster which will be tested using SSP-received power. A bus will be created to facilitate efficient power collection and transfer. Key design features will include: (A) maximization of receiving capability, (B) maximization of pointing, (C) reuse of structural elements for fuel storage

Spacecraft Bus

Launch system Orbit Ground System Communications Architecture

To be determined by launch provider Minor difference from transmitting spacecraft UND-developed Direct communications between UND/satellite

A slightly different orbit will be selected to facilitate a gradual separation to allow the characterization of system performance over a range of distances. A SDR-based ground station is under development at UND The communications will operate on the amateur frequency ranges, using an experimentalclass license.

on the receiving craft) are utilized to set and maintain the distance between the two craft. The craft will be maneuvered to multiple distances from the primary craft and the orientation between the two will be varied to facilitate characterization of the S2S-MWPT between the two. Each distance will be maintained for a period of time (which will vary based on the power cycle of the spacecraft) to facilitate multiple tests of the transmission at that distance and at each tested orientation at the distance. 4. Spacecraft systems overview This section describes the power transmitting and receiving spacecraft, including their deployment mechanisms and their power transmission and reception technologies. 4.1. Power transmitting and receiving spacecraft The proposed mission utilizes two satellites: an EELV Secondary Payload Adapter (ESPA)/Small Satellite (SmallSat) class spacecraft and a 1U CubeSat for the S2S-MWPT experimental mission. The 1U CubeSat is contained inside the SmallSat during launch. Upon separation from the

launch vehicle the following sequence occurs: (1) the 1U CubeSat is ejected from the SmallSat; (2) the 1U CubeSat deploys its rectenna array as (3) the SmallSat deploys its solar array. Fig. 1 depicts this sequence. Once this is accomplished, the MWP transmission and reception experiments can commence. The ESPA/SmallSat class spacecraft's primary purpose is to generate large amounts of electrical power from solar energy, convert this electrical power into an electromagnetic wave and transmit it into space towards the power receiving 1U CubeSat, as shown in Fig. 2. The electromagnetic wave travels to the targeted rectenna array on the power receiving 1U CubeSat where it is intercepted and converted into usable direct current power that is used to run all onboard subsystems, including the thrusters which thrusters ensure the 1U CubeSat maintains an appropriate distance from the transmitting SmallSat. Maintaining the spacecraft's position within minimum and maximum distance range bounds is absolutely vital for the S2SMWPT experiment, in order to obtain the appropriate power reception level needed to run all subsystems on the receiving spacecraft. Quantitative distance bounds between the satellites are determined using Eq. (8) and

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Fig. 1. On-orbit launch of sub-satellite and deployment of solar panels (graphical elements from [55]).

power that is acceptable to receive. For the purposes of experimental data collection accuracy, a higher level of pointing accuracy will be incorporated for this test mission than may be required on an actual service provider craft. Clark, et al. [56] have discussed the capability of creating an attitude determination and control system (ADCS) with a pointing accuracy of better than 0.11. This should be more than sufficient for mission needs. The selection of an actual ADCS unit will be made based upon a more extensive price versus accuracy trade-study. As the two satellites are in the sunlight portion of their orbit the MWP transmission experiments will be implemented. The power transmitting SmallSat will be generating electrical power from its solar array. Calculations for this are provided in Table 3. Using the process depicted in Table 1, the solar array area and mass were calculated as follows [57]: ððP e T e =X e Þ þðP d T d =X d ÞÞ Td ððð30WÞð36 minÞ=0:6Þ þ ðð300WÞð64 minÞ=0:8ÞÞ ¼ 64 min ¼ 403:125W

P sa ¼

ð1Þ

P BOL ¼ P o I d cos θ

ð2Þ

Ld ¼ ð1  DÞL

ð3Þ

P EOL ¼ P BOL Ld

ð4Þ

Fig. 2. Transmission of Microwave Power Beam [55].

Fig. 4, as well as the power requirements of the powerreceiving craft. In addition, the efficiency of the power reception and management/distribution units must be accounted for. With all of these details, the maximum separation that can support full power transfer efficiency can be determined. The required pointing accuracy is a function of the distance between the spacecraft and the minimum amount of

These calculations serve to size the solar array on the transmitting spacecraft and ensure that it is capable of producing the requisite power for the experiments. Fig. 3 shows solar array output power based on initial assumptions as outlined in the Table 3. The gimbal design [58] allows maximum sunlight to be collected for longer lengths of time as the transmitting satellite goes around the orbit.

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Table 3 Calculation Steps for Mass and Solar Array Area (based on [57]). Step

Reference

Nano SSPS

Amount of power that must be produced by the solar arrays, Psa

[57], pg. 650 (Orbit period  Te)

Pe ¼30 Wa, Pd ¼300 Wa Te ¼ 36 minb Td ¼64 minb A peak power tracking (PPTc) regulation scheme is assumed Xe ¼ 0.6 and Xd ¼ 0.8 Psa ¼403.125 W GaInP2/GaAs/Ge solar cells Po ¼ 387.144 W/m2 e Id ¼0.72 (Nominal value) θ  Sun Incident Anglef PBOL ¼ see Eq. (2) Performance degradation is 0.5% per year Ld PEOL ¼see Eq. (4) Asa E 1.7 m2

[57] Eq. (1) d Select type of solar cell and estimate power output Po, UTJ: Po ¼0.283  1368 W/ with the Sun normal to the surface of the cells m2 ¼ 387.144 W/m2 Beginning-of-life (BOL) power production capability, PBOL, Table 21-14 from SMAD 4 per unit area of the array Eq. (2) End-of-life (EOL) power production capability, PEOL, for Performance degradation the solar array UTJ: 0.5% per yrg Eq. (3) Eq. (4) Eq. (5) (3 year mission) Estimate the solar array area, Asa, required to produce the necessary power, Psa, based on PEOL an alternate approach Eq. (6) (28.3% efficiency) Estimate the mass of the solar array Eq. (7)

Asa E 2 m2 Msa E16.125 kg

%

3 years is used for the demonstration test mission assumption in this calculation. a The 300 W is broken up into 200 W of transmitted power, 70 W of assumed efficiency losses and 30Whr of battery storage. b From [57], pg. 650: “LEO spacecraft encounter at most one eclipse period each orbit or about 15 eclipse periods per day, with maximum shadowing of approximately 36 min.” Assuming a 100 min. orbit [8] this leaves 64 min daylight period. The batteries must charge and discharge about 5000 times each year giving an average depth-of-discharge (DOD) of  30%. SMAD 4 pg. 652 the DOD for Li-Ion in LEO is 20–40%. c From [57], pg. 653, PPT, the solar array and bus can operate at different voltages, and the PPT is functionally between the solar array and the bus, processing all the solar array power. d Spectrolab's Ultra Triple Junction (UTJ) GaInP2/GaAs/Ge solar cells with a bare-cell efficiency of 28.3%. The power input value for a planar solar array is the 1368 W/m2 or the solar constant (the amount of energy received at the top of the Earth's atmosphere on a surface oriented perpendicular to the Sun's rays). e From [57], pg. 645; Inherent Degradation, Id accounts for an assembled solar array being less efficient than the single cells due to design inefficiencies, shadowing and temperature variations, collectively. f From [57], pg. 647, the Sun incidence angle, θ between the vector normal to the surface of the array and the Sun line. g According to [57], pg. 644, performance degradation of a Triple Junction is 0.5% per year.

300

Solar Array Output

Power Output (W)

250 200 150 100

Power beginning of Life 3 Years 5 Years

50

1 7 13 19 25 31 37 43 49 55 61 67 73 79 85 91 97 103 109 115 121 127 133 139 145 151 157

0

Sun Incident Angle (deg) Fig. 3. Beginning of Life (BOL) Power and calculated solar panel degradation over 3 year and 5 years [59]. Calculations are made using a gimbal with 70 degrees of freedom.

4.2. Antenna size and satellite distance for power transfer optimization The electrical energy coming from the solar array is managed and distributed to the transmitting satellite subsystems. Power is channeled to the MWPT hardware and instantly converted from DC-to-RF and transmitted

into free space as an electromagnetic wave towards the power receiving craft. For this article the transmitted power was assumed to be 200 W, however, the amount of power density received at the center of the receiving location can be defined from Brown and Eves' [60] Eq. (3) as: P d ¼ At P t =λ D2 ¼ π d P t =4λ D2 2

2

2

ð5Þ

where Pd is the power density at the center of the receiving location, Pt is the total radiated power from the transmitter, At is the total area of the transmitting antenna, λ is the wavelength, D is the separation between the apertures, and d is the diameter of the parabolic reflector, if it is circular Estimated sizes for the transmitting parabolic dish and receiving aperture for this experimental mission are provided in Table 4. The transmitter uses the area of a circle while the receiver uses a rectangular area. From these estimated sizes and using Fig. 4 below the power density seen at across the rectenna is also shown in Table 4. From setting a desired value for the power density at the interface of the receiving array and from knowing the effective area (antenna aperture) of the receiver unit, the received power (assuming a uniform field) per element can be determined: P r ¼ P d  Aef f

ð6Þ

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Diameter of parabolic reflector (m)

Aperture size of receiver (m2)

Power Density (mW/cm2)

0.1 0.2 0.4 0.8 1 1.2

0.09 0.12 0.15 0.2 0.25

347.22 260.41 208.33 156.25 125.00

Ar=0.12

Ar=0.15

Ar=0.2

Ar=0.25

16 14 12 10 8 6 4 2 0 0

0.5

1

1.5

2

2.5

3

3.5

4

4.5

5

Transmitting Aperature Circular Area (m2)

Fig. 5. Shows the optimal distance required to achieve greatest power density transfer between the aperture antennas. The operating frequency is 10-GHz. Ar=0.09

Ar=0.12

Ar=0.15

Ar=0.2

Ar=0.25

9 8 7 6 5 4 3 2 1 0

Ar=0.09

0

0.5

1

1.5 2 2.5 3 3.5 Transmitting Aperature Circular Area (m2)

4

4.5

5

Fig. 4. Shows the optimal distance required to achieve greatest power density transfer between the aperture antennas. The operating frequency is 5.8-GHz.

The above equation can be used to estimate the power received per antenna element in a planar array assuming a uniform field distribution. Knowing the received power per antenna element and its impedance, the voltage can be determined: pffiffiffiffiffiffiffiffiffiffiffi V ¼ P r Rin ð7Þ Knowing the voltage input is important as the remaining portion of the circuit (as seen in Fig. 7) sees this as an AC voltage source thus the rectifying circuitry can be designed around an expected range of values. By knowing the power requirements at the rectenna array the RF-to-DC components and rectenna array architecture can be designed in order to accommodate this expected range of values. This procedure allows calibration of the power density at the receiving craft so that the components are capable to handle the power received. The receiver architecture and components are designed to the best of their ability to be able to handle a certain power level. This may require the power transmitted to be reduced or the distance between the satellites to increase which ultimately decrease tau in Eq. (8) to less than 100%. A balance must be reached to achieve greatest efficiency. Brown and Eves' [60] Eq. (1), and Fig. 2 provide a general equation for the aperture-to-aperture transmission efficiency in order to achieve optimum power density transfer based on a Gaussian profile scenario pffiffiffiffiffiffiffiffiffiffi τ ¼ At Ar =λD ð8Þ where At is the transmitter aperture area, Ar is the receiver aperture area, λ is the wavelength of the microwave power being transmitted, and D is the separation distance between the two apertures Fig. 2 of Brown and Eves is used to get a sense of the distance between the two spacecraft with the associated

Distance to receivier (m)

Distance to receiver (m)

Ar=0.09 Distance to receiver (m)

Table 4 Estimated diameter size of parabolic reflector and Aperture size of rectenna array.

199

Ar=0.12

Ar=0.15

Ar=0.2

Ar=0.25

350 300 250 250 200 150 100 50 0 0

0.5

1

1.5 2 2.5 3 3.5 Transmitting Aperature Circular Area (m2)

4

4.5

5

Fig. 6. Shows the optimal distance required to achieve greatest power density transfer between the aperture antennas. The operating frequency is 220-GHz.

power density and transmission efficiency based on a Gaussian profile scenario. Using their figure, τ was selected as 2.5 which indicates a 100% power density transfer efficiency. Then solving the equation above for D and using preliminary aperture size estimations from the Table 4, theoretical distance between the apertures to acquire 100% is determined and is shown in Fig. 4. Fig. 4 fixes five cases of the receive apertures and varies the size of the transmitter area to provide the optimal distance required to achieve greatest power density transfer between the two apertures. This data helps in setting the distance limits between the two spacecraft in order to supply enough power to the receiver to run its various subsystems. Similarly, Figs. 5 and 6 show results for 10-GHz and 220-GHz frequencies all other values and procedures are the same. Notice that the higher the frequency the greater the distance can be achieved for power density transfer as well as the option for smaller apertures. However, the disadvantage for higher frequencies is a decrease in component efficiency as well as increased atmospheric losses. 4.3. Transmitting and reception systems The aperture area of the power receiving spacecraft forms from the unfolding of panels that contain the rectenna array architecture. The transmitting craft will use the bottom and partial sides of its body to deploy into a main reflector dish. The arm holding the sub-reflector will ever so gracefully position itself in the appropriate position and lock in place. A corrugated conical horn antenna will be used as a feed mechanism. The transmitting antenna system is equipped with duel reflector Cassegrain architecture. Furthermore, as the arm that

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Fig. 7. Power flow diagram (revised from [55]).

holds the sub-reflector is extending into position the power receiving craft will be guided out of the transmitting craft into the space environment. The receiving craft will have an array of circularly polarized patch antenna elements. The Electrical Power System (EPS) is the major subsystem for both the power transmitting and power receiving satellites. Fig. 7 shows a flow chart for the MWP transmission and reception systems. As the Satellites enter daylight, the transmitting craft generates electrical power from its solar array, converts it into radio frequency (RF) energy (by means of a solid state power amplifier). The RF energy is transmitted via a corrugated horn antenna to a sub-reflector to the main reflector and then propagates to the rectenna array target on the power receiving craft located a distance away. The rectenna intercepts and converts the RF energy into usable dc power to run all subsystems. The end-to-end efficiency of the proposed system can be found in Table 5. Antenna misalignments between the two spacecraft were not included in the table. There are three scenarios to consider here (1) the two antennas maintain normal to the traveling wave but they vary in distance, (2) the antennas can misalign through rotation of either one of the satellite, and (3) either satellite can move out of point to point transmission exchange path. By incorporating a retrodirective beam mechanism on the two satellites will correct for most of these problems. The first problem can be viewed as a decrease in efficiency from the tau value mentioned above. In order to maintain 100% aperture to aperture efficiency three design options include: (a) increase or decrease the power level transmitted, (b) use thrusters to maintain the range of distance, and (c) use the buck-boost converter to regulate the voltage to the load. Part (a) is adjusted so that the received power to the components is not beyond their limit as it would destroy them. Part (b) uses the thrusters to balance and maintain the proper distance between the satellites. The second problem can be solved by equipping the receiving spacecraft with an active gimbal type system incorporated with a retrodirective array as well as using the thrusters to maintain balance.

Table 5 End-to-end efficiency estimations similar to figure in [67]. Segment

Efficiency

Notes

DC-to-RF Antenna

85% 70%

RF Collection Area Rectenna Elements Total

90% 80% 43%

SSPA [8] Dual reflector with corrugated conical horn Function of rectenna array size [8,26,68]

The third problem can be solved by incorporating a retrodirective array system to ensure the satellites remain in efficient power transfer exchange path. The design goal is to maintain the electromagnetic wave incident normal to the power receiving system. The proposed system is pursuing a hybrid antenna (HA) structure [61,62] with a Cassegrain reflector system (conventional offset structure) with a conical corrugated feed. The sub-reflector will have mechanically steerable capability for horizontal and vertical movement. Patch antennas for a retrodirective array system will be integrated on the back side of the subreflector that receive the pilot signal from the receiving spacecraft. That signal is then transferred to the motor controls of the subreflector for beam steering (self-phasing). The large aperture of the main reflector provides high-gain performance, the rotating subreflector provides beam-steering capability, and the retrodirective array system provides autonomous beam steering (self-phasing) and position sensing. Some papers have already began to explore the retrodirective array system [63–66]. The proposed work looks at advancing the HA structure and further expanding upon the ISY-METS experiment by incorporating a retrodirective array technology for a S2S-MWPT Experimental Mission. A limited backup power system will be incorporated on the receiving spacecraft in the form of placing solar panels below the storage location of the rectenna panels. This will result in the CubeSat having limited power generation

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capabilities in most orientations. This system will support troubleshooting and data retrieval from the CubeSat in the event of a transmission, catastrophic alignment or receiving system failure. A hybrid gimbaled rectenna and solar panel system may be able to be efficiently created; the design and development of this system will serve as a subject for future work. This type of hybrid receiver would potentially form an excellent backup solution for SSP utility consumer craft.

Human Controller

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Primary Spacecraft

Commence Test for each test

5. Launch The proposed mission will fly as a secondary payload onboard a commercial or government launch. As an ESPA/ SmallSat-class payload, it can be easily integrated onto rockets from several providers and into several locations on the rocket (at least eight [69], for example, on ULA vehicles like the Atlas IV). Several emerging launch providers may also be able to provide a dedicated or peershared launch capability for this type of mission at a more affordable rate [50,70]. A launch adapter plate is being utilized to facilitate delaying launch integration decisions until later in the mission planning process (once the specific rocket type is known). The satellite will connect to this plate and this plate will interface with the launch vehicle. Typically, the primary payload is launched first and, after the rocket has moved some distance from the primary payload, the secondary payloads are launched. The exact launch order amongst the secondary payloads or the order on a peer-shared mission would be negotiated with the launch provider. 6. Mission operations Mission operations commence with the ejection of the primary spacecraft from the launch vehicle. Once ejected, the spacecraft will wait a period of time (as specified by the launch provider) to allow it to move away from the launch vehicle. At this point, the sub-satellite will be ejected from the primary satellite and the primary satellite's solar panels will be deployed. With the two spacecraft in their operating configuration, the primary mission will commence. This will consist of transmitting power from the primary spacecraft to the secondary spacecraft. The secondary spacecraft will track the amount of power received and transmit this data back to the primary spacecraft. This experiment will be conducted at a variety of distances to characterize the performance at these distances. It will also be conducted at a variety of incident angles to characterize performance at both normal and non-normal angles of incidence. Fig. 8 summarizes the mission operations plan for the microwave power transfer portion of the mission. The power received by the sub-satellite will be utilized to operate its internal systems. It will also serve to power a set of onboard micro-cathode arc thrusters [29]. This will allow the performance of these thrusters to be characterized. These thrusters will be utilized to maneuver the secondary spacecraft to different distances. This will facilitate the power transfer experiment as well as allowing the

Secondary Spacecraft

Set Distance Confirm

Position Spacecraft

Notify Commence Acknowledge for each transfer Power Transfer

Mission Data

Test Results Analyze Results

Fig. 8. Mission Operations.

performance of the thrusters and their control system to be characterized. The performance of both (microwave power transfer and thrusters) will be characterized at different altitudes as the two spacecraft's orbits degrade towards atmospheric reentry. Two key variables will serve to define mission life. Orbital decay will place a maximum cap on the mission, which is a function of the altitude that the spacecraft are launched at, each spacecraft's ballistic coefficient and the current level of solar activity. Depending on the orbit utilized (as a secondary payload, the orbital options available to the mission will be limited), the onboard fuel may also serve to limit the mission. A number of factors will lead to fuel consumption which cannot be predicted a priori. The establishment of the initial configuration between the two craft (particularly if spinning must be arrested) will consume an unknown amount of fuel. The level of fuel required for the primary mission operations (i.e., moving from distance to distance to perform data collection) can be predicted. However, the level of fuel that will be required for orbital maintenance is unknown, in the absence of a target orbit. With a high enough altitude, this may not be required at all during the primary mission's duration. 7. Mission evaluation The main objective of mission evaluation is to complete successful demonstrations that could mitigate risks, buy down costs and advance technology relevant to future SSPS systems. There are several key experiments to consider.

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Demonstrating and validating key SSPS hardware elements and transmission characteristics. System performances for conversion and transmission efficiencies will be monitored. This will include a thermal analysis of the overall transmitter system onboard the bus platform. In addition characterization of thruster performance will be monitored. Ultimately, the thrusters are responsible to ensure the two spacecraft maintain a range of distance as described above and possible alignment/orientation during power beaming (i.e. beam riding). Other important data comes from analyzing system reliability, performances of components under severe thermal shocks, effects on component degradation over time in space, comparisons between land and space experiments. All these measurements will provide information for simulating the space environment on Earth for future component and system testing.

8. Conclusions and future work This paper has presented an overview of a near-term test mission to demonstrate and evaluate the efficacy of microwave power transmission as a mechanism to supply the power needs of other orbital spacecraft. It has provided an overview of the mission objectives and architecture. Multiple aspects of the mission have been analyzed to demonstrate its prospective feasibility and determine critical parameters for sizing the two mission spacecraft. Work on the design and development of the two spacecraft is ongoing. Structural and high-level thermal analysis has been largely completed. The electrical design of the power transmission and receiving system is a topic of ongoing work. Future work includes the completion of these designs, spacecraft construction, launch and mission operations.

Acknowledgments Work on the development of the Space Solar Power Small Satellite Test Mission at the University of North (Grant No. EPS-0814442) Dakota has been supported by numerous individuals including Robert Bernaciak, James Casler, Ben Kading, Michael Mann, John McClure, Sima Noghanian, David Poppke, Hossein Salehfar, Subin Shahukhal, David Whalen and Karl Williams. Facilities and equipment used in this project have been provided by the Department of Electrical Engineering, the Department of Mechanical Engineering, the Department of Space Studies, the Department of Computer Science and the College of Engineering and Mines. Limited financial support has been provided by the SunSat Design Competition. Thanks are also given to Paul Jaffe at the Naval Research Laboratory, Samudra Haque at George Washington University and Gerald Szatkowski at United Launch Alliance for providing information. We also thank Dr. Yen Lee Loh in the University of North Dakota Physics Department for his assistance regarding the physics behind Gaussian beams.

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