Transcript
Ben Blakeley Email:
[email protected] Phone: 435.881.9792
May 3, 2013
Dr. Donald Cripps Electrical and Computer Engineering Department Utah State University Dr. Cripps, Included is the final report for my senior project for your review. This project has been done in conjunction with Rincon Research Corporation and the rest of the ECE 5240 class. Rincon has identified a desire and a need to better understand the spatial variability of plasma in the midlatitude ionosphere and how this plasma alters signals as they pass through the mid-latitude ionosphere in space communications. My senior project has been to work with the rest of the ECE 5240 class and come up with a mission outline and CubeSat conceptual design. I have been working very closely with the rest of the ECE 5240 class and Dr. Swenson on this project, as well as Sid Henderson from Rincon Research. The class consists of four students, including myself. I have previously taken a CubeSat design course from Dr. Baker, as well as a Space Systems Engineering course from Dr. Swenson, and I believe these courses proved invaluable in the completion of this senior project. Both classes have helped me to develop a systems engineering mindset, which was absolutely necessary for a project of this type. I appreciate your consideration of this project. I also appreciate the time and effort you put in as a professor on behalf of the ECE department students. Sincerely,
Ben Blakeley
Senior Project Final Report for
RaPTIR (Radio-wave Propagation Through Ionospheric Regions)
ECE 4850 May 3, 2013
Ben Blakeley
Instructor Approval _________________________________________ Don Cripps Department of Electrical and Computer Engineering Utah State University
_________________ Date
Instructor Approval _________________________________________ Charles Swenson Department of Electrical and Computer Engineering
_________________ Date
Utah State University
Abstract The ionosphere is a layer of the atmosphere that spans from about 85 km up to 600 km. The ionosphere plays an important role in radio-wave propagation. Lower frequency waves are deflected, and higher frequency waves are degraded in terms of amplitude, phase and polarization. The fact that lower frequency waves are deflected is a key factor to modern day communications. The characteristics of this layer are due largely to solar conditions. The varying solar conditions cause great variability at different daily times, seasons, position and also altitudes. Different techniques have been used to measure the ionosphere in the past, but they are limited in spatial coverage and resolution. The RaPTIR mission presents a concept for a CubeSat constellation which collects simultaneous measurements from different spatial positions. With this detailed information it will be possible to better understand ionospheric physics and improve associated models. The mission concept of operations, required scientific payload, and the host spacecraft are presented in this report.
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Table of Contents Abstract …………………………………………………………………………………………... I List of Figures ...……………………………………………………………………………….... IV List of Tables ...………………………………………………………………………………….. V Acknowledgments …………………………………………………………………………….... VI 1. Introduction ……………………………………………...………………………………….… 1 1.1 Problem Statement ………………………………….……………………………………. 5 1.2 Design Summary ………………………………….……………………………………… 5 2. Mission Overview ……………….……………………………………………………………. 6 2.1 Mission Objectives ……….………………………………………………………………. 6 2.2 Science Overview ………….…………………………………………………………….. 7 2.3 Mission Concept of Operations …….……………………………………………………. 8 2.4 Payload …………………….……………………………………………………………. 10 2.4.1 Payload Antenna …….………………………………………………………….... 10 2.4.2 Payload Concept of Operations ……….…………………………………………. 13 2.4.3 Payload Receiver and Processing ……….……………………………………….. 15 3. Spacecraft Bus and Ground Segment …….…………………………………………………. 17 3.1 Propulsion Subsystem …………….…………………………………………………….. 18 3.1.1 Propulsion Concept of Operations ……………………………………………….. 21 3.2 Command and Data Handling Subsystem …….…………………………………...…… 21 3.3 Communications Subsystem ……………….………………………………………….... 22 3.4 ADCS Subsystem ……………………….……………………………………………… 24 3.5 Power Subsystem ……………………….…………………………………………….… 24 II
3.6 Mechanical Subsystem ………….………………………………………………………. 25 3.7 Mechanisms .…………………….………………………………………………………. 26 3.8 Thermal Subsystem …....………….…………………………………………………….. 27 3.9 Ground Segment …....……………….………………………………………………….. 27 4. Final Scope of Work …………………………….…………………………………………... 28 5. Cost Estimation ………………………………….…………………………………………... 29 6. Project Management Summary ………………………….…………………………………... 29 7. Conclusion ……………………………………….………………………………………….. 31 Appendix A: Bibliography ……………………………………………………………………... 33 Appendix B: Resume …………………………………………………………………………... 34
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List of Figures Figure 1: Ionospheric layers (daytime vs. nighttime) …………………………………………… 2 Figure 2: Radio-wave Propagation in Ionosphere ………………………………………………. 3 Figure 3: Mission Concept of Operations ……………………………………………………….. 8 Figure 4: Payload Concept of Operations ……………………………………………………… 13 Figure 5: Payload Block Diagram Flow Chart ………………………………………………… 14 Figure 6: Cross-Section View of RaPTIR ……………………………………………………... 17 Figure 7: ΔV Required for Extending Satellite Orbit at Different Altitudes …………………... 20 Figure 8: Cadet-U Radio ……………………………………………………………………….. 22 Figure 9: RaPTIR Product Breakdown ………………………………………………………… 28 Figure 10: Gantt Chart …………………………………………………………………………. 31
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List of Tables Table 1: Payload Antennas Trade-Off …………………………………………………………. 11 Table 2: Telemetry Link Budget ……………………………………………………………….. 23 Table 3: Power Budget …………………………………………………………………………. 25 Table 4: Mass Budget ………………………………………………………………………….. 26
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Acknowledgments I would like to take this opportunity to thank the rest of the ECE 5240 class who are not a part of the ECE 4850 class this semester. Without them this project would never have made it as far as it has. My contributions to the project were a relatively small part of the overall project. The other members of the class conducted critical research and came up with information that was crucial to the development of the mission. I am grateful for their contributions, and also for their willingness to participate in my design review presentation. I would also like to thank Dr. Charles Swenson for his guidance and assistance in the design process throughout the semester. Even though he had to travel doing consulting for NASA for much of the first part of the semester, he was always available for help. One day as a class we even went over some information with him on the phone for the project as he was sitting in an airport in Boston waiting to board a plane to come home for the weekend. Without his assistance, as a class we never would have made it to this point in the design. I would also like to thank Stephanie Sullivan of the Space Dynamics Laboratory for stepping in and conducting class for Dr. Swenson and helping to keep things moving along in his absence. I would also like to thank Dr. Sid Henderson of Rincon Research Corporation for agreeing to be the ‘customer’ for this project and for giving us a foundation of requirements to base our design decisions on in order to complete this conceptual design. I would like to thank the Space Dynamics Laboratory for assisting in generating 3d images of our CubeSat design. Finally, I would like to thank Dr. Cripps and Laura Vernon for their assistance, support, and words of encouragement. It is thanks to this caliber of educators and professors who are so concerned with the success of their students that the engineers of the future are making such a big difference. VI
1.
Introduction For recent decades, man has endeavored to travel into space in an attempt to better
understand what lies beyond the surface of earth. But up until the launch of the unmanned Sputnik I in 1957, man, nor anything of his creation, had not previously ventured into space. Throughout man’s exploration of space, he has encountered that the higher above the surface of the earth you go, the more conditions change from what we are used to on the surface. These changes in the atmosphere as elevation increases have been divided into certain areas and given corresponding names. The main levels of the atmosphere as we know them today are the Troposphere (7-16 km above earth’s surface), Stratosphere (16-50 km above earth’s surface), Mesosphere (50-80 km above earth’s surface), and Thermosphere (80-500 km above earth’s surface). As these regions are traversed upwards from the surface of the earth, various changes occur. Temperature decreases as you go up in the troposphere, then increases as you go up in the stratosphere; temperature again decreases as you go up in the mesosphere, and then increases as you go up in the thermosphere. But, apart from the temperature increasing, as we increase in altitude we encounter various physical phenomena such as nacreous clouds, the ozone layer, noctilucent clouds, meteors, and polar lights. Many of these things can affect space travel as well as space communication. From the upper part of the mesosphere up to the top of the thermosphere is a region called the ionosphere. The ionosphere is characterized by its high number of ionized particles produced by solar ultraviolet radiation. The ionosphere can be divided into sub-layers including the D, E, and F layers, and during the day the F layer divides further into the F1 and F2 layer and the D layer disappears, as can be seen in Figure 1.
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Figure 1: Ionospheric layers (daytime vs. nighttime) The ionosphere suffers variations during large periods of time, which are due to solar cycles and seasonal time. The electron density is higher during solar maxima, and typically during the summer. In addition to temporal variations, there exist spatial variations across the earth. Because solar radiation varies with latitude, four distinctive regions can be observed: polar, auroral, midlatitudes, and equatorial. There are, however, certain anomalies and perturbations to this simplified model. These include the winter anomaly, equatorial anomaly, and ionospheric storms. Ionospheric storms are characterized by a large increase in the electron density caused by solar flares or Coronal Mass Ejections (CME). In addition to these external sources, energy and momentum enters the ionosphere from below through tides, planetary waves, and atmospheric buoyancy waves
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generated in the lower atmosphere. The significance of these perturbations and anomalies differ depending on the ionospheric region. The ionosphere has a huge impact on radio wave propagation. Radio waves with low frequencies will be reflected at different altitudes, depending on their frequency and angle of incidence. This phenomenon has been used for many decades for long range transmission of radio signals. Higher frequency radio waves, such as those used in satellite communications, are able to pass through the ionosphere. They do, however, suffer from attenuations, polarization change, and diffraction. This effect can be seen in Figure 2.
Figure 2: Radio-wave Propagation in Ionosphere Critical frequencies have been identified as the maximum frequency that will be reflected for a perpendicularly incident incoming wave. The critical frequency at the F2 layer electron density peak, or foF2, is the maximum signal frequency the ionosphere is capable of reflecting for a wave with a perpendicular angle of incidence. The foF2 is located at an altitude called the height of maximum F2, or hmF2. 3
The electron density in the ionosphere has been studied for over 50 years using different methods. These methods can be divided into two main groups: ground observations and in-situ observations. Ground observations have been performed with ionosondes and incoherent scattering. An ionosonde is the emitting of radio-waves at different frequencies in the vertical direction and then measuring the reflected signal. In-situ observations are based on rocket or satellite instruments. Instruments on-board rockets are capable of obtaining an accurate profile of the electron density at different altitudes, but can only collect data for a few minutes and with a limited spatial profile. Some conclusions have been drawn that the amount of plasma in the ionosphere from space weather is what affects communication signals, and a better understanding as to how and why space weather affects them will help us improve space communications. Plasma makes up roughly 99% of all matter in the universe, but we don’t see it on earth because it is too cold down here for most matter to reach that state. The sun is made of plasma, as are all the stars. Their intense heat can turn nearby gas to plasma. Knowing this, one can see how a spacecraft would encounter a large amount of plasma on orbit. If we were able to better understand why the ionosphere affects High Frequency (HF) radio waves in space communications, this could help us to improve space communications as well as to better understand the ionosphere and space weather. A better understanding of space weather could lead to numerous advances in fields that are affected by space weather, such as agriculture, weather forecasting, and communications, to name but a few. This project report summarizes a CubeSat mission design that will help us better understand the plasma densities in the mid-latitude ionosphere, as well as how this layer of the atmosphere affects communications the way it does.
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1.1
Problem Statement Rincon Research Corporation LLC has identified a need to better understand why the
ionosphere affects HF radio-wave transmission, and they have asked the ECE 5240 class to come up with a conceptual CubeSat design to accomplish this task.
1.2
Design Summary This report outlines the conceptual design of a CubeSat mission which will perform
measurements on signals that have been transmitted through the ionosphere and compare the signals received to the original signals. The received signals will be originally transmitted via CODAR, WWV, and AM radio transmitters. The key characteristics of the signals that will be monitored for change are the amplitude and phase. A better understanding of how much these two characteristics of a signal change, and why, will help us to better understand the ionosphere and space weather.
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2.
Mission Overview The goal of the RaPTIR mission is to better understand the spatial variability of plasma
densities in the mid-latitude ionosphere during transition periods and over scale sizes that affect HF radio-wave propagation.
2.1
Mission Objectives The main objectives were defined by Rincon Research after input from the ECE 5240
class. These objectives, which have driven the mission concept and design, are as follows: •
Observe the variability of 10 to 1000 km density structures that exist over the range of 200 to 400 km in altitude during four transitional local time intervals: Midnight (00000100), Sunrise (~0600-0700), Noon (1200-1300), and Sunset (~1800-1900), all at magnetic mid-latitudes (30°-60°).
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Simultaneously observe the magnitude and phase of HF (1-20 MHz) ground transmitters from a space-based constellation of three HF sounding receivers. The ground stations of interest are CODAR, WWV, and AM radio signals.
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Collect between 2 and 8 minutes of continuous observations simultaneously from at least two satellites over ground transmitters and retrieve this data for study within 14 days.
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Collect at least 25 data samples for each of the time intervals over at least a 3 month season (with a goal of 6 months) and during an equinox and a solstice.
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2.2
Science Overview The different layers of the ionosphere and their corresponding levels of ionization allow
for signals of varying frequency to pass through or be reflected back towards the Earth. By measuring the radio signals during a complete season, the current model of the ionosphere could be refined. Above the hmF2, the foF2 will be obtained by measuring the received signals in the 1-20 MHz band and calculating the cut-off frequency. The received signals will be compared with the known transmitted frequency to determine attenuation and delay characteristics induced by the ionosphere. The night-time critical frequency is typically 2-4 MHz and the daytime critical frequency is typically 10-15 MHz. This mission will take measurements during quiet ionospheric hours (noon and midnight) as well as the transition hours (sunrise and sunset), to understand diurnal variations. When the satellites’ altitude decays below the hmF2, the constellation will be able to receive low frequency signals such as AM radio stations, which emit frequencies of 1-1.5 MHz. These signals will be affected by the lower layers of the ionosphere (D and E), and therefore it will be possible to obtain information about their effect on low frequency (LF) wave transmission. The major contribution of this mission with respect to past measurements is the ability to measure small scale ionospheric structure with controlled resolution. Data from the RaPTIR mission will help to better understand the spatial density of plasma in the ionosphere.
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2.3
Mission Concept of Operations
Figure 3: Mission Concept of Operations The mission Concept of Operations can be seen in Figure 3. A constellation of three CubeSats will be transported to the International Space Station (ISS) on one of the regularly scheduled resupply launches. The satellites will be released from the J-SSOD deployment system on the ISS. Their orbits will be approximately 380-400 km, slightly lower than that of the ISS. The satellites will be released with slightly different ΔV to achieve an initial separation. The J-SSOD CubeSat release module on the ISS can provide appropriate release velocities. The CubeSats will have a circular decaying orbit, covering the complete altitude range of interest during their lifetime. In order to extend the lifetime of the spacecraft at low orbit and to control satellite spacing, a propulsion system will be implemented.
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A constellation of three CubeSats provides greater variation for differential measurements by allowing one CubeSat to drift further away from the other two. It will also increase reliability by providing redundancy, because in the event of one spacecraft failure, the other two can still fulfill the mission. The preferred signal transmitters will be the Coastal Ocean Dynamics Applications Radar (CODAR). CODAR stations are primarily used to measure ocean currents and tides, but have also been used to measure the height of the ionosphere. CODAR stations are the preferred ground transmitters for several reasons: •
They are widely dispersed across US coastal regions, allowing for considerable satellite contact time.
•
They emit a chirp with a very narrow bandwidth (18-100 kHz).
•
They are monopole antennas, with a transmitted power of 40-50 W. The power density at the satellites location will be enough to obtain an acceptable Signal-to-Noise Ratio (SNR).
•
The frequency range of operation for CODAR is 3-50 MHz, with many stations operating within the band of interest (1-20 MHz). Other signals that will be received are WWV signals (not an acronym, but a name) and
AM radio waves. WWV signals are continuously transmitted timing signals at frequencies of 2.5, 5, 10, 15, and 20 MHz out of Fort Collins, Colorado. WWV signals are controlled by local atomic clocks traceable to the United States National Institute of Standards and Technology’s (NIST) primary standard by GPS common view observation. Once the constellation orbit has decayed below the hmF2, the system will also receive AM radio signals.
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Access time for three ground signal transmitters was simulated in the Satellite Tool Kit (STK) software, obtaining more than 40 valid accesses for each time period. There are 79 CODAR stations within the desired frequency range, and therefore the contact time to collect 25 measurements per time period can reasonably be achieved with a great amount of margin. CODAR, WWV and AM radio signals are transmitted using monopoles, and consequently have a vertical polarization. However, the ionosphere rotates the signal polarization. The satellite receiver will be capable of receiving both linearly and circularly polarized signals.
2.4
Payload The RaPTIR payload is a custom design of a tunable radio that will allow ground station
users to select certain frequency bands of interest from averaged spectra snapshots of received data. The payload has been designed to fulfill the science collection objectives of the mission.
2.4.1 Payload Antenna The payload antenna had two very distinctive design drivers: a very low frequency band, and a very high bandwidth. It was necessary to calculate the range of wavelengths that could be received based off of the mission objectives. The equation λ = c/f was used to calculate the different wavelengths, with c being the speed of light (3x108 m/s), and f being the frequency of the wave (in hertz). The band of interest includes wavelengths ranging from 15 to 300 meters, and thus any antenna
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on a small spacecraft will be electrically short. Using small antennas implies a trade-off between dimensions, gain, and bandwidth. Multiple antenna solutions were studied (see Table 1). While the Moebius antenna has a very high bandwidth and high gain compared to the others, it is considerably difficult to deploy a one meter diameter loop. The magnetic rod antenna is very easy to deploy because of its reduced dimensions, but its poor gain at the HF band resulted in an impractical solution.
Table 1: Payload Antennas Trade-Off Parameter
Moebius
Magnetic rod
Short monopole
Short dipole
Bandwidth
0.1-30 MHz
VLF - 3 MHz
Outside resonance
Outside resonance
Polarization
Linear
Linear
Linear
Linear
Gain
-30 dB
-50 dB
-80 dB (without matching)
-80 dB (without matching)
Radiation pattern
Omnidirectional Omnidirectional Omnidirectional Omnidirectional
Dimensions
1 m diameter
Advantages
High bandwidth Easy to deploy Moderate gain
Less weight and Symmetrical dimensions than radiation dipole pattern and drag
Disadvantages
Very difficult to Low gain at HF deploy
Low gain Asymmetrical drag and radiation pattern
10 to 15 cm length
1 m length
2 m length
Low gain
The antenna selected for the RaPTIR mission consists of two crossed dipoles, with arms 1 meter in length. The main reason is their relative ease of deployment. They also have an omnidirectional radiation pattern which reduces spacecraft attitude control requirements. The use 11
of dipoles over monopoles is based on obtaining symmetrical drag, and a uniform radiation pattern. Dipoles have linear polarization, and thus, to cover both linear polarizations two independent receivers will be implemented. Although this requires double power and data, it records the polarization information of the incoming radio wave(s). The use of small dipole antennas presents two problems: a very low effective length and very high capacitive impedance. To overcome the impedance matching problem, two solutions were evaluated: 1. Impedance matching network: this is a common solution for high SNR, but with a limited bandwidth. 2.
High Impedance Amplifier (HIA): by using higher input impedance than the antenna impedance, reasonable gains in the complete band can be obtained. Because the payload required sampling the complete frequency band, the HIA was
chosen. A careful design with low input noise can be used to amplify the signal without degradation (assuming the external noise is dominant). In this way, the problem of low effective length was overcome. The detailed design of the HIA has been identified as one of the main design drivers in order to obtain acceptable instrument performances in terms of dynamic range and sensitivity. Another intrinsic advantage of the HIA is a constant gain over the entire bandwidth. Because the HIA is sensible only to the open circuit voltage (proportional to the electric field), the Friis equation becomes independent of the signal wavelength.
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2.4.2 Payload Concept of Operations Initial science objectives were to receive and process all signals from 1-20 MHz, store the raw data in on-board memory, and downlink the full data on the next overpass of a ground station. With the limited amount of available ground stations and contact time, the effective downlink rate would not be enough to transmit all the raw data measured. The current Concept of Operations is based on selecting sub-bands at certain frequencies, where the signals of interest are located, and downlinking only these sub-bands. This considerably reduces the amount of data necessary to be downlinked, because all the signals of scientific interest have a narrow bandwidth compared to the full receiver bandwidth. In order to select these sub-bands, ‘snapshots’ of the complete bandwidth will be received regularly, and downlinked. The science radio Concept of Operations can be seen in Figure 4.
Figure 4: Payload Concept of Operations 13
The sequence for the payload will be as follows: 1. The ground station operator will send operation times for the science radio to the spacecraft. 2. The spacecraft will collect data at the specified times, storing the complete bandwidth in mass-memory. 3. The spacecraft will generate a spectrum from small snapshots of the raw data. 4. During the next ground station overpass, data snapshots will be downlinked. 5. Upon review of these snapshots, the ground station operators will determine the bands of interest and upload them to the spacecraft. 6. The bands of interest will be processed and downlinked on the next overpass of a ground station.
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2.4.3 Payload Receiver and Processing The basic flow of the payload, along with its components, can be seen in Figure 5.
Figure 5: Payload Block Diagram Flow Chart (Flow included for only one of two polarizations; each antenna will have its own receiver) Signals from the active antennas will be acquired by a homodyne receiver (a homodyne receiver combines two waves of identical frequency; the received signals will have the same frequencies, but different polarizations). The signal will be filtered for the band of interest, 1-20 MHz, in order to avoid saturation of the amplifier (the frequency band of interest for the mission is a very large bandwidth and the amount of signals and noise should be reduced as much as possible). The filtered signal will be amplified and conditioned to suitable levels for the ADC. The ADC will have a resolution of 16 bits in order to provide the necessary dynamic range. This will allow handling of high power signals such as AM radio (between altitudes of 200-250 km), as well as lower power CODAR signals. The sampling rate for raw data will be 45 Msamples/s. 15
The raw data will be allocated in a mass-storage device with a time and position tag obtained from the ADCS subsystem. Processing of the data will be completed by the On-Board Computer (OBC). Spectra of the raw data will be generated, averaged, and sent to ground for review. The processor will generate a 1 ms spectrum of raw data for 10 consecutive intervals, and the resulting 10 spectra will be averaged. This sequence will be repeated for every 30 seconds of raw data. Once all raw data has been processed in this way, the averaged snapshots will be sent to ground for review. Averaging the spectrum data will allow reduction of noise and spurious signals. The phase will be discarded in the snapshot telemetry data to reduce the downlink rate required. Upon review of the downlinked spectrum averages, ground users will select certain subbands of interest. The selected sub-bands will be digitally shifted to base-band by use of a Digital Down-Converter (DDC). A DDC converts a digitized signal centered at some intermediate frequency to a base-banded signal centered at zero frequency, and decimates the remaining data. This makes it possible for the signal of interest to be shifted to baseband, and once the remaining unwanted data is decimated, the signal of interest has a much narrower bandwidth than it previously had. This means there will be much less data to transmit. The baseband bandwidth will be configurable according to the bandwidth of the signal of interest and the final resolution reduced to optimize the downlink rate. It was decided to use compression algorithms to reduce the data snapshots as well as the user-selected bands of interest. The use of lossy compression algorithms was agreed to, and they will allow a high compression ratio without compromising the science requirements.
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3.
Spacecraft Bus and Ground Segment For the RaPTIR design, the PEARL platform, which was developed at the Space
Dynamics Laboratory (SDL), was chosen (see Figure 6). The main reasons for this decision were that it fulfilled all spacecraft requirements, and because SDL has a long history of developing and delivering electronics that have consistently met and exceeded expectations. The PEARL platform includes an electronics stack with five boards: ADCS sensor interface board, bus interface controller board, electrical power supply board, battery board, and one board reserved for the payload. It also includes a modular structure and a customizable number of solar panels.
Figure 6: Cross-Section View of RaPTIR The RaPTIR constellation will be 3-axis stabilized, in order to minimize drag and increase mission lifetime. Another important reason for 3-axis stabilization is so that the receiving antennas are constantly pointing at the earth. Two types of stabilization were examined: 3-axis stabilized and Spin stabilized. The main drawback of Spin-stabilization for the 17
RaPTIR mission would be that the spacecraft would be constantly spinning and thus the antennas would only be pointing toward earth a fraction of the time. Since ground station contact time is already so limited with only two available ground stations, the choice was made to make the satellite 3-axis stabilized. The PEARL bus provides an ADCS system capable of a pointing accuracy of 1°.
3.1
Propulsion Subsystem The required spatial separation of the RaPTIR satellites is between 10 and 1000 km, with
measurements being made uniformly throughout this range. These measurements must also be made throughout an altitude range of 200 to 400 km, requiring RaPTIR to operate in very low Earth orbit and face significant drag from the upper atmosphere. Because of these two factors, a propulsion system will be included on the RaPTIR bus. Busek’s Micro Pulsed Plasma Thruster (μPPT) system was chosen for this mission for a number of reasons, including low system volume, relatively low power (6 W peak power), and, most importantly, high total impulse (1200 N-s). The μPPT system uses high voltage to cause an arc that ejects solid Teflon propellant to provide thrust to the spacecraft. The μPPT system is supplied with the necessary high voltage generator and control electronics, providing a simple interface to the CubeSat processor. RaPTIR will be using the μPPT system exclusively for propulsion, so each thruster (48 in total) will be pointed in the anti-ram direction of the spacecraft. The propulsion system will perform two functions: 1. Satellite separation control 2. Orbit maintenance to counteract drag 18
Satellite separation control can be achieved by instructing one of the satellites to fire its thrusters in the direction of the velocity vector and move to a slightly higher orbit. The other satellites remain in the lower, original orbits with smaller radii, which allow them to move ahead of the satellite that performed the burn. Once the satellites are separated as desired, the forward satellites perform an identical burn and their rate of separation reduces to nearly zero. This method of satellite separation control can be achieved quite economically according to studies performed using STK’s Astrogator orbit design tool. Two identical satellites with ballistic coefficients of 105 kg/m2 were placed in a circular orbit with an altitude just below the ISS. When the spacecraft perform burns, their relative ground track velocities change. An example of this is outlined below: 1. RaPTIR2 performs an initial burn and moves above RaPTIR1. RaPTIR1 now has a smaller orbit radius, allowing it to move faster than RaPTIR2. 2. RaPTIR1 performs an identical burn 5 days later, and the constellation’s rate of separation decreases significantly. 3. After 5 more days, RaPTIR1 fires its thrusters, allowing RaPTIR2 to have a faster ground track velocity. 4. When the satellites meet, RaPTIR2 performs a burn and the two satellites fly in close formation. Each burn requires only 10 cm/s of ΔV, so many of these separation control maneuvers could be performed with very little propellant. Additionally, minimal active attitude control is necessary because the lawn dart configuration helps to ensure that the thrusters will fire in the ram direction (forward direction), as simulated.
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The majority of the propellant will be used for orbit maintenance at low altitudes. Measurements below 300km altitudes are of great scientific interest, and therefore it is important to extend the lifetime of RaPTIR in this range. Atmospheric drag increases dramatically below 300km, so the propulsion system will have to be used often (15% of the time at 250 km; up to 50% of the time at 200 km). The ΔV required to offset the force of drag increases dramatically with decreasing altitude, as can be seen in Figure 7. These values were calculated for a satellite with a 105 kg/m2 ballistic coefficient maintaining its orbit for 36 days. The calculations use the NRLMSISE 2000 atmospheric model, and assume higher than average solar activity.
Figure 7: ΔV Required for Extending Satellite Orbit at Different Altitudes
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3.1.1 Propulsion System Concept of Operations The propulsion system Concept of Operations while above 300km are as follows: 1. Ground control will determine the desired arrangement of the RaPTIR constellation. 2. Satellite tracking data and position telemetry will be used to determine RaPTIR constellation separation and calculate the necessary maneuvers. 3. Ground control will instruct each satellite to perform burns at specific times to achieve the desired separation. For altitudes above 300 km, propulsion should not be needed for orbit maintenance. According to STK’s Lifetime Tool, a RaPTIR satellite will take about 175 days to naturally decay from 390 km to 300 km. RaPTIR will take full advantage of this by taking as many measurements above 300 km as it can without performing orbit makeup burns.
3.2
Command and Data Handling Subsystem The processor inside the PEARL stack is a LEON 3FT capable of executing up to 60
million instructions per second at 75 MHz. Studies performed show that this level of processor performance is well above what will be required for the RaPTIR mission. The following processes will be controlled by the LEON 3FT processor on the PEARL stack: •
Process the spectrum analysis of the data snapshots.
•
DDC processing of the sub-bands of interest selected by ground stations.
•
Payload control, switching on/off the electronics, and triggering the data acquisition.
•
ADCS collection and data and processing to obtain the position and orientation.
•
Propulsion system control (length of the burn) 21
•
Mechanisms deployment
•
Ground station communications
3.3
Communications Subsystem The communication system is based on the Cadet-U radio, pictured in Figure 8. The
Cadet-U radio was developed by L-3 Systems, in collaboration with SDL. The Cadet-U, which was developed for DICE (a previous project at SDL), provides a raw 3 Mbit/s downlink, a 19.2 kbits/s uplink rate, and data storage.
Figure 8: Cadet-U Radio This highly integrated subsystem includes both transmitter and receiver functionality in a single unit, operating in half-duplex mode. The 3 Mbit/s downlink uses QPSK modulation and
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Forward Error Correction (FEC) coding. The use of FEC limits the effective bitrate to 2.6 Mbits/s. The Cadet-U radio is used with four monopole antennas phased from the same driver, obtaining a quasi-isotropic radiation pattern. The telemetry link budget can be seen in Table 2. From these calculations, it is evident that there exists a positive telemetry margin; this signifies that the system will be capable of collecting more science data than required.
Table 2: Telemetry Link Budget Design element
Value
Total Data To Downlink
Units
339 Mb/overpass
Overpasses per Day
2 Overpass
Science Data per Day
677 Mb/day
Downlink Telemetry Rate
2.6 Mb/s
Packet Overhead
7% Unitless
Ground Contact Time
3% Unitless
Single Satellite Allocation Time
50% Unitless
Ground Station Availability
24% Unitless
Downlink Capability
796 Mb/s 18% Unitless
Telemetry Margin
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3.4
ADCS Subsystem The requirements for the ADCS subsystem are moderate pointing accuracy for
propulsion maneuvers and antenna pointing, and a high position and pointing knowledge accuracy for science data timestamp and physical position. The ADCS will be composed of a GPS receiver, a patch antenna for the GPS, a star sensor, multiple magnetometers, three torque coils and a reaction wheel. With these instruments, a position knowledge accuracy of better than 10 meters and orientation knowledge accuracy better than 70 arc-seconds can be obtained. The ADCS will maintain the minimum drag orientation of the CubeSat within 1°.
3.5
Power Subsystem RaPTIR’s power system will utilize a combination of solar panels and batteries. The
solar panels system will consist of 3 body-mounted panels and three deployed panels in a lawn dart configuration. The nadir (earth-facing) panel will be excluded (see Figure 6). Based on calculations performed in Excel, this will provide an orbit average of about 15 W, which is sufficient for RaPTIR’s subsystems, as seen in the Power Budget in Table 3. Table 3 includes all the major subsystems and reflects worst case conditions (such as the heavy usage of the propulsion system at the end of the mission).
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Table 3: Power Budget Peak Power (mW)
Subsystem Payload
Avg. Power (mW)
700
14
Propulsion
6,000
3,300
C&DH
4,006
2,505
Communications
7,022
561
ADCS
2,485
1,001
Power
13
12
1,200
0.01
10
3
21,436
7,396
Mechanisms Thermal Total
Total factoring in efficiency and 10% margin
10,576
The batteries used will be lithium-polymer cells. Four cells in series will be needed to provide the required energy storage while RaPTIR is in eclipse.
3.6
Mechanical Subsystem (Structure) Because this mission will be launched from the ISS using the J-SSOD deployer, a
CubeSat was a natural choice for the structure of this spacecraft. The PEARL platform includes a three-section modular structure to aid in assembly, integration, and testing. This bus also has a solid plate on each base to give the structure a greater shielding against radiation. The mass budget of the complete satellite, divided into subsystems, is presented in Table 4.
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Table 4: Mass Budget Subsystem
Total mass (g)
Payload
400
Propulsion
547
ADCS
301
Communication
240
Mechanisms
21
C&DH
502
Thermal
51
Power
479
Structure
86 3,113
Total mass
10%
Margin Total mass + margin
3.7
3,424
Mechanisms The bus and the payload have multiple mechanisms. After release from the ISS, and
waiting a preset period of time to ensure sufficient separation from the ISS, solar panels, CadetU antennas, and the payload antennas will deploy. While solar panels and Cadet-U antennas will use a conventional spring released mechanism, the payload antennas required special consideration due to their length. The proposed solution consists of four Storable Tubular Extendible Member (STEM) antenna mechanisms. These antennas are formed by specially treated metals (steel or beryllium copper), rolled and compacted in small dimensions. To deploy, a small motor or pyrotechnic is actuated, and the spring energy releases. Once deployed the antennas take a long and rigid shape.
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3.8
Thermal Subsystem The thermal subsystem of the spacecraft is principally passive. The components used
have wide operational temperature ranges, and only the battery will require active thermal control. A heater mounted to the battery housing will be operated with a bimetal thermostatic switch. The payload receiver performances will drift with temperature. A stable temperature is required to avoid large gradients (in particular for the HIA). It is also to be noted that the receiver noise will increase as the temperature increases. Thermal design must assure a high dissipation from the receiver board to the structure.
3.9
Ground Segment The proposed location for the ground station for the RaPTIR mission is Wallops Island,
VA because of the large 18 meter dish antenna used for this link. A secondary dish to be built in Logan, UT has also been proposed to help downlink the large amount of data anticipated for collection on this mission. These ground stations will communicate with the spacecraft using a modem, which was developed by L-3 Systems for the DICE mission. The ground station will provide the Ultra High Frequency (UHF) feed to the ground station modem. The modem will be connected to a RaPTIR program-supplied computer, which will serve as the primary mission operations center.
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4.
Final Scope of Work Up to this point, as a class we have come up with design solutions to many of the large
problems that a mission such as this would encounter. We broke the mission down into its main subsystems, as can be seen in Figure 9, and did research and trade studies, individually as well as collectively, to determine the best solution for each respective system. No definitive hardware has been specified (beyond the PEARL stack, CADET-U radio, and μPPTs), and no software algorithms have been written. These would be the next steps in the design process.
Figure 9: RaPTIR Product Breakdown 28
5.
Cost Estimation Throughout the semester, the RaPTIR mission has been an unfunded conceptual design.
Dr. Sid Henderson from Rincon Research Corporation stepped in at the beginning of the semester at the behest of Dr. Swenson to act as a customer for this design, giving the class a feel for what it is like to work in the field of engineering and have to interface with customers and complete a design to meet certain requirements. To Dr. Henderson and the science community in general the information that a mission like this could potentially yield, were it to receive fullfunding and be carried out through implementation and launch, is of great interest. The requirements that Dr. Henderson initially gave the class had been well thought-out and tailored so that the mission would produce precise results. He was not just doing a favor to Dr. Swenson and stepping in blindly. Having said that, the concept of the mission and our design could potentially receive funding. After the semester is over, as a class we will give a presentation on our design to Dr. Henderson and potentially others from Rincon Research, as well as to engineers and scientists at SDL.
6.
Project Management Summary With the help of Dr. Swenson, we divided up the following responsibilities between the
four members of the class as follows: Ben: •
Data flow and payload design.
•
On-orbit collection capability and necessary memory capacity.
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Kashele: •
Telemetry and link budget.
Chris: •
Orbit characteristics and collection times
•
Lifetime of the CubeSats and how the constellation will evolve and change over time.
Julio: •
Concept of receiving antenna on spacecraft.
•
Determine which type of antenna pattern will best meet the mission needs.
The remaining subsystems not mentioned in the previous bullet points were designed collectively by the students in the class to the best of our ability using software tools to generate as accurate of simulations and numbers/quantities as possible. All of the members of our class are at the senior/graduate student level in our studies and we relied on knowledge and tools obtained in previous classes, such as Dr. Swenson’s Spacecraft Engineering class, to make the design decisions that were made for RaPTIR.
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1/7 1/14 1/21 1/28 2/4 2/11 2/18 2/25 3/4 3/11 3/18 3/25 4/1
4/8 4/15 4/22 4/29 5/6 5/13
Determine mission objectives Divide tasks amongst team Run trade studies/do research Refine objectives Re-evaluate payload design Explore alternatives Define remaining subsystems Create 3D mock-up of CubeSat w/SDL Prepare for Rincon presentation Bring subsystems together Final report rough draft Final report final draft
Figure 10: Gantt chart A Gantt chart for activities carried out throughout the semester can be seen above in Figure 10.
7.
Conclusion With the RaPTIR system, it is possible to obtain useful information about the ionosphere
by making measurements in different temporal conditions and with different and controlled spatial resolutions. These kind of measurements are not possible with big satellite constellations because their location is normally driven by their host spacecraft (like the GPS), and therefore the separation between satellites cannot be controlled for science purposes. In the mission outlined in this document, a very inexpensive constellation can perform this task with great flexibility.
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The RaPTIR mission will provide valuable information regarding the spatial distribution of plasma in the ionosphere. By measuring the bending of radio waves at different points in the ionosphere and how the density of plasma at those same points affects radio wave transmission, we will gain a better overall knowledge of how this layer of the atmosphere affects signal transmission and communication in general. Although the design has been performed with a small constellation of three satellites in mind, it is easily scalable to a bigger constellation. This would improve the spatial information and therefore yield a more accurate model. In the same way, a smaller constellation of two satellites would be possible without significant loss of science information, with the advantage of improving the total downlink per satellite.
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Appendix A: Bibliography Figure 1: Ionospheric layers (daytime vs. nighttime) www.britannica.com Figure 2: Radio-wave Propagation in Ionosphere www.nasa.gov
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Interior Alaska Roofing Summer of 2009, 2010, 2011, and 2012 • I worked as a roofer and learned the process of flat-roofing, and also did some metal work, shingling work, and learned minor carpentry skills. I worked 50-60 hour weeks and was given lots of responsibilities in spite of my arguably small amount of experience. I was relied upon for my dependability, reliability, and hard work ethic. Inovar, Inc. September 2011 to March 2012 • I worked as a test engineer intern. I worked with the test engineers to troubleshoot and fix various products, as well as design and implement test benches. Companies that I worked on products for included L3 communications, Varian Medical, and Cobb Tuning. Space Dynamics Laboratory October 2012 to Present • I work as part of a research team of engineers and we are working with a custom design of a fiber filter spectrometer. We are in the process of collecting data and analyzing it. Our goal is to find out how to exploit the data in order to make this particular model of the spectrometer a viable choice in industry and in terms of hyperspectral research.
ADDITIONAL QUALIFICATIONS In my time studying Electrical Engineering at USU I have taken courses on and feel that I am proficient in the Following areas/programs: • Signal processing (including some exposure to DSP) • Analog circuit analysis and debugging • Digital circuit design and analysis • Matlab • C++ programming • Assembly language programming • Cadence design suite • Verilog HDL • Use and programming of microcontrollers • Electromagnetics • Antenna theory and design • Exposure to optical practices in an Optical Engineering course • STK (Satellite Tool Kit) – I am a certified user of the program • Microsoft Office Suite (extensive use of Excel) Currently I am working on my senior project in a space systems design course. We are working on a conceptual design of a CubeSat in conjunction with Rincon Research Corporation in Arizona that will measure how signals are altered as they pass through the ionosphere in space communication. I was in the Boy Scouts of America and at the age of 18 became an Eagle Scout. I lived in northern Peru for two years doing humanitarian service and am fluent in Spanish. I enjoy challenging tasks and I thoroughly enjoy anything that has to do with the outdoors, especially camping and hunting. I am an avid runner and enjoy playing basketball. I am a very dependable and responsible person.
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BEN BLAKELEY BEN BLAKELEY
RECENT JOB EXPERIENCE
[Type your address] [Type your phone number] [Type your e-mail address]
Utah State University Completed 4.5 years Studying Electrical Engineering. I am currently a senior and will graduate in the spring of 2013.
356 North 300 West #6 Logan, Utah 84321 435.881.9792
[email protected]
EDUCATION
