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Purdue Imaging Satellite PI-SAT

Giles Goetz Marcos Hasebe John Hawkins Chris Patterson

Ramses Ramirez Aric Simmons

Colin Sipe

School of Aeronautics and Astronautics Purdue University

May 3, 2002

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Executive Summary The school of Aeronautics and Astronautics department at Purdue

University has developed a senior and graduate level class to design a low-cost micro

satellite. The Purdue University – Imaging Satellite (PI-SAT) is designed to take

pictures of Purdue University from a sun-synchronous orbit. The satellite is

approximately 50 Kg, and it is a cube of 0.5 meters on each side. The satellite is

gravity gradient stabilized, and is equipped with two CMOS cameras. Spacecraft

power will be supplied by Li-Ion batteries and body mounted solar panels. The

spacecraft will be capable of downloading at least one picture per day. The

spacecraft structure has been designed for launch on the Ariane 5 ASAP program.

The goal is to have a functioning imaging satellite communicating with

Purdue University’s ground station for at least three months, at a cost of less than 2

million dollars. An operational PI-SAT could be completed within a couple of

years. Completion of PI-SAT will enhance the reputation of Purdue University as a

leader in the Aerospace community.

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Table of Contents

Section: Page:

3. Mission Statement………………………………………………..….4

4. Mission Objective…………………………………………………....4

5. Satellite Description………………………………………………....4

6. Concept of Operations………………………………………………

7. Major Design Requirements……………………………………..…6

8. Orbit Selection…………………………………………………...…..8

9. Launch Vehicle Integration………………………………….…….11

10. Spacecraft as a System………………………………………….….13

11. Payload……………………………………………………….……..17

12. Attitude, Determination & Control……………………….………19

13. Communications……………………………………………..……..26

14. Command and Data Handling……………………………….……34

15. Power…………………………………………………………….….40

16. Thermal……………………………………………………………..

17. Structures and Mechanisms………………………………….……43

18. Summary of Satellite…………………………………………….…50

19. Appendix……………………………………………………..……

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3. Mission Statement We will make a low cost student-built satellite to observe the Earth’s atmosphere and land masses so that we can post pictures on Purdue’s website for public relations purposes.

4. Mission Objective

To take pictures of the Earth from orbit and send image data back to ground stations.

5 Satellite Description

Figures 5.1, and 5.2 show an unexploded and exploded view of the satellite, respectively. Figure 5.3 shows the axis system used for the satellite. These pictures are with the antennas compact.

Figure 5.1. Unexploded view of the satellite.

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Figure 5.2. Exploded view of the satellite.

Figure 5.3. View with axis system.

The total mass of the satellite is 50.4 kg. The breakdown of the mass budget can be found in Section 10.3.

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7 Major Design Requirements 7.1 Customer Attributes The satellite customer is Purdue University. The University’s requirements are:

• PI-Sat shall be a small low cost satellite • PI-Sat shall have the ability to take pictures of the Earth from on orbit and transfer

those pictures back to campus. 7.2 Engineering Requirements Broken down by subsystem Payload (Camera) The PI-Sat camera shall:

• Have a resolution of at least 1 km • Require less than 10 Watts of power

Orbit Selection

• Pi-Sat shall be in an orbit that allows for regular transfer of data back to the ground station at Purdue.

Requirements from Launch Vehicle In order to meet the requirements of the Ariane Structure for Auxiliary Payload (ASAP) program, PI-SAT, in its launch configuration, shall:

• Weigh under 120 kg • Have a cross section of under 0.6x06 m and a height of less than 0.71 m • Have moments of inertia of less than 20 kgm2 • Have a center of mass located within 5 cm of the center of area in the x-y plane

and less than 45 cm away from the mounting plate in the z direction • Have a maximum longitudinal acceleration of –7.5/+ 5.5 g and maximum lateral

acceleration of +/- 6g Structures and Mechanisms The PI-Sat structure shall:

• Support Pi-Sat’s weight through launch and orbit operations. • Structure natural frequency greater than 90 Hz. • Be simple to build and relatively inexpensive. • Use no exotic materials • Be capable of being built by students.

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Power PI-SAT’s power system shall:

• To supply all subsystems with power for life of mission • Support a lifetime of at least a year

Attitude Determination and Control PI-SAT’s attitude shall be such that:

• The camera shall be pointed towards the Earth • Attitude can be determined and controlled with enough accuracy to place the

desired object of a picture within the field of view of the camera Communications PI-SAT’s telecommunications system shall:

• Handle different types of data, such as attitude, command and housekeeping data. • Have the capacity, and signal strength to transmit image data. • Bit error less than 1%

Command and Data Handling PI-SAT’s CDH system shall:

• Interpret flight plan • Store data (OS, image files, housekeeping log) • Control power modes of all components • Control ADCS to orient PiSAT autonomously • Transfer data between memory and communications • Take temperature measurements • Check power • Control communications to send data and receive commands from ground station

Thermal

• PiSAT’s thermal system shall keep electronics and batteries within their operating ranges

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8 Orbit Selection 8.1 Orbital Parameters/ Keplerian Elements

Our satellite was chosen to behave in a near polar sun synchronous orbit with the following Keplerian Elements

Orbital Type Sun-

Synchronous Altitude (km) 800 Inclination 98.7 degrees Mean Motion 14.3 Orbital Period 100 minutes Eccentricity 0 (circular orbit)

Figure 8.1. Orbital characteristics.

Using the orbital parameters given in Table I, the Keplerian Elements of three consecutive passes were monitored in order to calculate the RAN times and confirm that for a sun-synchronous orbit, the satellite solar times would remain approximately constant.

UTC 14:28 Longitude(deg) -72.96 Latitude (deg) 0 RAN (hours) 4.84 Satellite Solar Time 9:36

Figure 8.2. Descending First Pass Keplerian Elements and Times.

UTC 16:04 Longitude(deg) -96.495Latitude (deg) 0 RAN (hours) 6.433 Satellite Solar time 9:38

Figure 8.3. Descending Second Pass Keplerian Elements and Times.

UTC 17:45 Longitude(deg) -123.15Latitude (deg) 0 RAN (hours) 8.21 Satellite Solar Time 9:32

Figure 8.4. Descending Third Pass Keplerian Elements and Times.

As expected for the three passes, the satellite solar time remains constant as would be true for a sun-synchronous orbit.

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The following UTC times refer to the times in which the satellite was ascending towards the equator for the above three scenarios:

First Pass 21:36 Second Pass 21:38 Third Pass 21:32

Figure 8.5. Ascending times.

Finally, the following table represents the time in UTC in which the satellite first makes contact with the ground station footprint for each pass, and the amount of time in which the satellite was in this footprint for data transmission (duration) respectively.

Crossover Time(UTC)Duration

(min) First pass 14:11 11:32 Second pass 15:50 11 Third pass 17:30 11:32

Figure 8.6. Footprint crossover and duration times.

8.2 Motivation for Orbit Selection/ How orbit meets requirements of s/c and mission It is expedient for a communication satellite receiving data from the same point

on earth with every pass to be able to design an orbit, which is conducive to such data transfers. Since a sun-synchronous orbit allows the craft to pass over the same point of earth at roughly the same time each day, it was expedient for our purposes. By being able to maintain regular contact with the Lafayette ground station, the number of pictures that can be taken can be maximized by a sun-synchronous orbit. In addition the generous altitude at 800km allows a large enough footprint to be able to receive large amounts of data with every pass while not being so high as to compromise resolution of the payload camera. 8.3 Trade Studies/ Comparisons

One possible trade study would be to vary the altitude of the satellite by making it lower, for instance. The tradeoff being made in this case is compromising data transfer time with camera resolution. Although the images will come out crisper with larger pixels, the average amount of time that the satellite will pass over the ground station will be less than for the main design. The following are the first and second pass footprint durations for low altitude case at 700km:

Altitude (km) 700Period (min) 98.8Mean Motion (rev/day) 14.579First pass (min) 11:30Second pass (min) 9:26

Figure 8.6. Footprint Crossover Duration Times for Altitude of 700km.

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As expected, due to a smaller footprint since the satellite has a smaller “field of view” with respect to the ground station, the amount of data transfer time would be somewhat less than for the main design. For the second pass, the data transfer time already decreased by over two minutes. Another minor point is that the altitude decrease leads to a lower orbit (higher mean motion) and a slightly shorter period.

To be complete, the opposite end of the spectrum had to be analyzed as well. If the altitude of the craft were to be increased, camera resolution would be compromised for a larger footprint with respect to the ground station. For this scenario, an altitude of 900 km was considered:

Altitude (km) 900Period (min) 102.9Mean Motion (rev/day) 13.9824FD first pass 14:24Second pass 16

Figure 8.7. Footprint Crossover Duration Times for Altitude of 900km. *NOTE: This orbit took 4 passes until the communications footprint could be “seen.”

Due to the footprint size increase, placing the satellite at this altitude would incur

a noticeable increase in data transfer time. However, since it was decided that a minimum pass of 10 min per orbit was suffice for the data transmission, it was thought that it would be unwise to further compromise camera resolution. For this reason, it was agreed to stick to the baseline design.

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9 Launch Vehicle Integration 9.1 Launch Vehicle Selection The Ariane 5 launch vehicle provided the best solution for our spacecraft. The Ariane Corporation has a program designed specifically for small satellites, which is termed Ariane Structure for Auxiliary Payload (ASAP). There was a very specific manual for ASAP on the Ariane Corporation web page, which provided detailed information about the payload envelope, adapter, and mass limits of the spacecraft. Excerpts from the manual can be found in the appendix. 9.2. Selection Process Atlas IIAS and Delta II were considered as launch vehicles for our spacecraft, but they did not offer the same benefits as the Ariane 5. The Ariane 5 had a very detailed payload manual for small satellites; the Atlas and Delta did not have specific programs for small satellites. The Ariane 5 also has small satellites launch on almost every vehicle because they have the ASAP, which is completely independent of the primary payload. The Atlas and Delta did not have a dedicated area on the launch vehicle for small satellites, and the payload envelope, adapter, and mass limits of the spacecraft would depend on the primary payload of the launch vehicle. The Ariane frequently delivers satellites in our desired orbit and altitude, reference section 8 for more information. The ASAP program can carry and deploy small and medium satellites on LEO, SSO, MEO or GTO orbits. 9.3 Requirements of Launch Vehicle The ASAP manual had very detailed requirements stated, which are found in Figures 9.1 –9.5. The mass of our spacecraft is only half of the maximum allowed on the ASAP for a micro satellite, therefore we were not as concerned about mass as we initially thought we would be. The spacecraft fits within the dimensions given, and the antenna will be at a sufficient angle as to stay within the boundaries. The moments of inertia are much less than the maximum of the ASAP, as shown in reference Figure 10.6. The center of mass of the spacecraft is within the constraints of the launch vehicle, as shown in reference Figure 10.4. The maximum acceleration of the spacecraft was modeled in ANSYS; reference Section 17.4 for results of the analysis.

Mass of Auxiliary Payloads Micro < 120 Kg Mini 120 < s/c < 300 Kg

Figure 9.1 - Maximum allowable mass of spacecraft.

Envelope of Auxiliary Payloads Cross Section 0.6 X 0.6 (m)

Height 0.71 (m) Figure 9.2 - Maximum allowable dimensions of spacecraft.

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Maximum Mass Moments of Inertia (Kg-m2)

Ixx < 20 Iyy < 20 Izz < 20

Figure 9.3 - Maximum allowable moments of inertia

Center of Mass (m) X < +/- 0.05 from

Center of Area Y < +/- 0.05 from

Center of Area Z < 0.45

Figure 9.4 - Maximum allowable location for center of gravity.

Maximum Acceleration Longitudinal -7.5 / +5.5 g

Lateral +/- 6 g Figure 9.5 - Maximum acceleration of spacecraft.

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10 Spacecraft As a System 10.1 Internal Layout

Once all the components of the spacecraft were known, the most efficient placement of the components could be achieved. After much iteration the following layout was decided upon, as seen in Figure 10.1. For an illustration of which subsystem each color component belongs to, refer to Figure 10.2. An advantage with this layout is that all components are close together. This limits the amount of wiring that will have to be used, and thereby decreases cable power loss.

Figure 10.1. Internal layout of the satellite.

Number Part1 Cameras2 Antenna3 Transmitter/Receiver4 Batteries5 Power Conditioner6 Nutation Dampener7 Torque Rods8 Magnetometer9 CDH Hardware

Figure 10.2. Identification for 10.1.

Other considerations for the layout were also stability issues. If a certain component was heavy, then another component that had a similar weight was placed in the same position, just on the opposite side of the satellite. By doing this, the satellite is becoming more axially symmetric, which by having this creates a more stable object.

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10.2 Component Packaging The advantages of this layout, is that no components will be “hanging” from the sheet during the launch phase. This allows the components to be attached to the sheets much simpler, and with less bulky fastening methods. The batteries are together and have aluminum along the sides of them to hold them in place. The components of the Command and Data Handling system are housed in an aluminum box. The reasoning behind this was to protect from radiation effects and also it was a way to enclose most of the electronics of the satellite as possible. 10.3 Mass budget of Spacecraft

The mass budget for the spacecraft is given in Figure 10.3. As shown in the figure, the total mass of the spacecraft is 50.5 kg. This is a very good mass, which should increase our chances of getting on the launch vehicle; reference Section 9 for more launch vehicle data. Figure 10.4 shows a pie chart of the mass budget, which illustrates how each subsystem, contributes to the overall mass of the spacecraft. The margin of 16.74% is based on a goal of weighing under 60 kg, so it is the difference between the actual mass of 50.5 and 60 kg.

Component Mass (kg) Component Mass (kg)

Camera 0.011 Power Hub 1.2Frame-Grabber 0.1 Computer 2

Battery 1 0.3 Gravity Boom 2.2Battery 2 0.3 Tip mass 1Battery 3 0.3 Thermal Blankets 1Battery 4 0.3 Solar Panel 1 1Battery 5 0.3 Solar Panel 2 1Battery 6 0.3 Solar Panel 3 1Battery 7 0.3 Solar Panel 4 1Battery 8 0.3 Magnetometer 0.2

Antenna 1 0.3 Magnetorquer 0.66Antenna 2 0.3 Sun Sensor 1 0.04Antenna 3 0.3 Sun Sensor 2 0.04Antenna 4 0.3 Beams 21.3Receiver 0.65 Skin 9.57

Transmitter 1 Adapter 1.89Total: 50.461

Figure 10.3. Mass budget of spacecraft.

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Mass by Group

9.82%

5.00%

3.51%

4.39%

16.74%

0.19%

57.47%

1.65%

1.23%

Payload

Power

CommunicationsCD&H

GravityBoomThermal

Attitude

Structure

margin

Figure 10.4. Pie chart of spacecraft mass.

The center of gravity of the spacecraft is in a very favorable location, which is

given in Figure 10.5, reference Figure 5.3 for location of coordinate system and dimensions. This location is very good because the spacecraft uses a gravity gradient system.

Location of Center of Mass (boom extended)

Axes Location (m) X 0.250 Y 0.242 Z 0.335

Figure 10.5. Location of Center of Mass.

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10.4 Mass Moments of Inertia

The moments of inertia with the beam stowed are given in Figure 10.6, the spacecraft will be in this configuration during launch. Figure 10.7 shows the mass moments of inertia for the spacecraft on orbit.

Mass Moments of Inertia (boom stowed) (Kg-m2)

Ixx 2.837 Iyy 2.680 Izz 2.620

Figure 10.6. Mass Moments of Inertia, stowed boom.

Mass Moments of Inertia (boom extended) (Kg-m2)

Ixx 13.654 Iyy 13.555 Izz 0.765

Figure 10.7. Mass Moments of Inertia, extended boom.

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11 Payload 11.1 Objective of Payload

The payload of the satellite consists of a pair of color CMOS digital cameras. The cameras are mounted on the Earth facing side of the satellite. The cameras will be used to take pictures of the ground and atmosphere from orbit. 11.2 Component Selection and Sizing

The specifics for the camera are in Figure 11.1, and a figure of the camera is in Figure 11.2. The camera was selected for its small size and power requirement as well as the wide field of view. The camera meets the requirements of the 1 km resolution as well.

Size 22mm(W) x 25mm (H) x 30mm (D)

Weight 11 grams Power 0.36 Watts (30mA @ 12VDC)

Operating Temperature -10 to +55 deg C

Format NTSC (510 x 492) pixels Lens 6.0 mm

Image Area 4.95 x 3.54 mm

Field Of View (FOV)

Length 47.3 deg Width 33.8 deg

Resolution

Length 1.37 km/pixel Width 0.988 km/pixel

File Size 734 Kilobytes

Figure 11.1: Specifications for Payload Camera

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Figure 11.2. Picture of CMOS Color Camera.

The small size of the CMOS camera makes it idea for use on a satellite. The only drawback is the camera is not space qualified. Overall the picture will be 700 km x 490 km in size from orbit. 11.3 Subsystems and Operations

The camera requires a power load of 0.36 Watts from the power system when taking a picture, the rest of the time the camera will be off. In order to take a picture the camera will be connected to the computer’s systems via a cable connected to the PCMCIA frame grabber card. The frame grabber card will take an image from the camera and send it to the computer for storage. The camera will be turned on just prior to taking a picture and then turned off once the picture has been taken.

11.4 Trade Studies Originally a CCD camera was chosen for the camera of the satellite, because many had been used on previous spacecraft. CCD cameras are larger and require 10+ Watts for operation. CMOS cameras were then suggested as an alternative because of their low mass and power consumption. Also originally a PCI frame grabber card was to be used as the interface with the computer, but the computer had no means of connecting to a PCI frame grabber card so an alternative was found. That information is located in section 9.2.1 of the report.

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12 Attitude Determination and Control Subsystem 12.1 Major Requirements Attitude determination shall be accomplished such that the pointing of the camera can be determined to an accuracy of less than 10 degrees for taking pictures. This is a rather low level of accuracy, but given that the camera has such a wide field of view, it is enough to ensure that the desired object of the picture is within view. The attitude control system shall provide the ability to stabilize the satellite in a nadir pointing orientation so that the camera can view desired ground targets. Again, control is only needed to maintain nadir pointing within 10 degrees. Note that this is only a 2-axis control requirement; the third axis (spin of the spacecraft) is unconstrained. As well, the subsystem must be capable of failure recovery. A failure is defined as occurring when the spacecraft pointing has deviated more than 10 degrees from nadir, the worst case being a 180 degree flip. It is not uncommon for a gravity gradient stabilized satellite to flip upside down and stabilize pointing away from nadir. 12.2 Strategy and Operations 12.2.1 Strategy To achieve nadir pointing, gravity gradient stabilization was chosen because it is the simplest method available and largely passive. This strategy requires a long boom with a tip mass deployed from the spacecraft, a passive nutation damper, and a small actuator that will be used to establish nadir pointing after spacecraft deployment and do failure recovery. A magnetorquer, consisting of a collection of torque rods, has been chosen as the actuator. The torque rods require sensors to sample the local magnetic field, so a magnetometer is also required. This can also be used for attitude determination and gives course 3-axis data. The primary devices used for determination, however, are sun sensors, which give more reliable 2-axis data. Two sensors are needed, one fore and one aft to give good coverage so the sun is always in sight of one of them. Their data is to be supplemented with magnetometer data. Sun sensors will be vital during times when the magnetorquer is active since the fields generated by the torque rods will interfere with magnetometer data making the magnetometer useless for attitude determination during these periods. Strategy choices are outlined in the table. Requirement Strategy

Nadir Pointing Gravity Gradient Stabilization Determination to < 10 o Sun Sensors supplemented with Magnetometer data

Control to < 10 o Magnetorquer and a passive nutation damper Failure Recovery Magnetorquer capable of countering worst case torques

Figure 12.1. Strategies to meet requirements.

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Figure 12.2

12.2.2 Operations There are four phases to the operations of ADCS. In chronological order they are: 1) Establish nadir pointing after spacecraft deployment This involves the use of sensors to determine the tip-off spin rate and the magnetorquer to counter this initial disturbance and align the spacecraft. This operation should be performed before the boom is deployed, since the boom would serve to increase the spacecraft moments of inertia, increasing the needed control effort. 2) Deploy Boom Once the spacecraft has established nadir pointing, and before disturbances have significantly altered the pointing, deploying the gravity gradient boom will establish the desired passive control. 3) Monitor Orientation with Sensors This allows the spacecraft to always have attitude determination and recognize when a failure has occurred, such as being flipped over with the wrong end pointing towards nadir, or if disturbances have turned the pointing too far away from the nominal. During this period, control is entirely passive and accomplished by the gravity gradient and the nutation damper. 4) Do Failure Recovery If failure has occurred then the magnetorquer is to be employed to recover the nadir pointing. Being able to accomplish this means having the ability to counter all possible disturbance torques as well as countering the worst-case gravity gradient torque itself in order to recover the nominal orientation from a complete failure with the spacecraft pointing 180 degrees away from nadir.

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Figure 12.3 Operations outline.

12.3 Component Selection and Sizing The Primary drivers for component selection have been first mass, then power, then cost. 12.3.1 Attitude Determination Components Several possible components were considered before deciding on the final configuration. Component Accuracy Applicability

Magnetometers ~ 5 o Required for field sensing for magnetorquer Useful during eclipse when sun sensors are offline

Earth Sensors 0.1 o Heavy and consume a lot of power Won’t need this level of accuracy anyway

Sun Sensors 0.5 o Very light and low power Good for use when magnetorquer is on and magnetometer will not be useful.

Gyroscopes 0.001 o/hr Must periodically reset reference position More complicated than necessary

Figure 12.4. Choice of components for determination

It was decided after much deliberation that Earth sensors would not be needed. The level of accuracy they afford is unnecessary and cannot justify the expense in mass or volume they take. See the trade study for further information. Gyroscopes, also, were discarded early on as unnecessary. Sun sensors and a magnetometer should be adequate for our purposes. Sun Sensors The chosen sun sensors are Model 0.5 Sun Sensors from Optical Energy

- ±0.5° 2 Accuracy over 100° field of view

- Weight < 40 g

- Power < 50 mW

- Reliability = >0.999 for a 15 year missions (longer than needed)

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The sun sensors were chosen for their low mass and power requirements, and small dimensions. Their accuracy is actually better than needed, but sun sensor accuracy is usually this good or better and this does not have any negative effects. Magnetometer The chosen magnetometer is a TFM100S Attitude Control Magnetometer from Billingsley Magnetics. - Course accuracy

- Weight = 200g - Power = 0.56 W - Radiation Shielded

The magnetometer was chosen again for small size and mass and low power. It also has the advantage of being space qualified and radiation shielded. 12.3.2 Attitude Control Components Several options were considered for attitude control. Momentum wheels and control gyros are typically very heavy and powerful, and so they were discarded as options. Using thrusters would have complicated the system by adding the need for a propulsion subsystem, including its extra mass. A yaw axis momentum wheel would have allowed yaw control, but this was found to be unnecessary. 2-axis passive gravity gradient control with nutation damping and no 3rd axis control is adequate. A magnetorquer was chosen as the actuator to be used for attitude establishment and failure recovery. Both the magnetorquer dipole and the gravity gradient boom sizes had to be considered to ensure that the system would work as desired. The main consideration for sizing the components (magnetorquer, gravity gradient boom, and tip mass) is to ensure that the control authority is larger than the disturbance torques. The largest disturbances come from solar pressure and the aerodynamic effect. The sizes of these effects are directly related to how long the boom is, since a longer boom increases the center of gravity to center of pressure distance (cg-cp). It is necessary to ensure that the gravity gradient torque is not overpowered by the disturbance torques. This drives the system towards having either a longer boom or a heavier tip mass to increase the gravity gradient torque. Increasing the boom length, however, also increases cg-cp, which increases the disturbances. At the same time, the magnetorquer must be sized to counter the worst case torque for failure recovery.

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Gravity Gradient Boom and Tip Mass It was decided to use an SSTL-Weitzmann deployable boom from Surrey Satellite Technology Ltd.

- Deployed Boom length: 3 m - Boom mass: 2.2 kg - Tip mass: 1 kg - Boom Diameter: 11 cm - Stowed length: 27 cm

There were few off-the-shelf gravity booms to choose from. This one has the advantage of being very compact when stowed and allowing the customer to select sizing. The boom and tip mass were sized to minimize cg-cp while still giving a strong gravity gradient torque. With this configuration, the spacecraft moments of inertia are:

Ix = 16.25 kgm2 roll Iy = 16.11 kgm2 pitch Iz = 0.80 kgm2 yaw

Which gives a maximum gravity gradient torque at 800 km altitude of: Tg = (2.474 x10-5)sin(2θ) Nm

Where θ is the nutation angle. Also, the cg-cp distance is less than 0.6 m which gives maximum worst case disturbance torques of:

Ts = 2.953x10-6 Nm Ta = 9.874x10-7 Nm

Note a gravity gradient torque has been achieved which is stronger than the disturbances for moderately small θ. Magnetorquer In the event of a failure, the magnetorquer must be sized to counter the worst case torques, both from disturbances and gravity gradient, this worst case torque is:

Tw=Ts+Ta+maximumTg = 2.87x10-5 Nm

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At 800 km altitude, the worst case magnetic field strength is B = 4.304x10-5 T, so

to generate a torque greater than the gravity torque requires a minimum magnetic dipole moment, Dmin, of: Dmin = Tw/B = 0.666 Am2 The chosen magnetorquer consists of three orthogonal TR1UPN Torqrods from Ithaco Inc. Torqrod Properties (each): Linear Dipole Moment: 1.1 Am2 (minimum) Mass: 100 g Power: 0.28 W With this dipole, the torqrods are capable of applying a torque of

T = B*D = 4.73x10-5 Nm This is more than enough to counter the disturbances and do failure recovery.

Nutation Damper The nutation damper consists of a viscous ring damper mounted in the spacecraft, perpendicular to the yaw axis. The viscous ring damper is a very simple device that uses fluid friction to dissipate energy in a nutating body. The ring consists of thin walled tubing, rigidly mounted to the spacecraft body, and partially filled with a viscous fluid. When the spacecraft turns about the spin axis (which is left uncontrolled), centripetal force causes the fluid to push against the inner wall of the tube. This creates a viscous friction force that opposes any motion of the fluid. When nutation occurs, the off-axis spin causes the fluid to move up and down. The viscous friction opposes this motion, dissipating the energy in the nutation. The sizing of the nutation damper is an area still under development. 12.4 Predicted Performance The attitude determination components are capable of giving the required 10o accuracy. The sun sensors alone actually give better accuracy, but for times when the sun sensors are not useful, such as in eclipse, the magnetometer can be used. The magnetometer alone can do the task, but cannot be used when the magnetorquer is in operation, so sun sensors are still required if for no other reason than to monitor. The satellite will be able to stay nadir-pointing within the required accuracy of 10o. Maximum total disturbance torques have been calculated to be on the order of 4x10-6 Nm in the worst case. At a nutation angle of 10 deg the gravity gradient torque is 8.5x10-6 Nm making it the dominant torque in the system, even in the worst case. So disturbance torques should not be able to force the satellite out of alignment once nadir pointing has been established. Furthermore, the nutation damper will be eliminating nutation oscillations.

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If failure does occur, however, the magnetorquer is capable of applying a torque which can counter both the disturbance torques and the worst case gravity gradient torque, so it will be strong enough to re-establish nadir pointing. Note that since the magnetorquer is capable of re-establishing nadir pointing, it must also be capable of initial pointing establishment after satellite deployment since that actually requires less torque. 12.5 Trade Studies Early in the design process it was thought that a pointing accuracy of better than 2o (or less) would be needed in order to point the camera at a desired target. At this point the design included a need for one or even two Earth sensors which could weigh as much as 5 kg in total and drain 5 W of power.

However when the camera choice was made and the chosen camera was found to have such a large field of view, it was realized that an accuracy of less than 10o would be adequate to place the target within the picture. So the requirement for high accuracy had been dropped and subsequently the Earth sensors were removed from the design. The current design is therefore far less accurate while still meeting the overall requirement of being able to take the desired picture, and saves a lot in mass and power.

Also it is common for gravity gradient stabilized spacecraft to have a yaw wheel for 3rd axis control. This was originally an element of our design, which was again heavy and consumed a lot of power. In reality though we have no requirement for 3rd axis control. 2-axis control is enough to keep nadir pointing. So the yaw wheel was also removed.

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13 Communication 13.1 Design goal and requirements The primary communication design goal is to enable communication between satellite and ground station. Communication needs to be bi-directional, enabling telecommands for changing the mode and parameters of the satellite, telemetry for receiving housekeeping information and handling image data. The telecommunication characteristics will be decided by the requirements from payload and orbit characteristics. The orbit is expected to be polar, with an inclination of 98 degrees, and altitude of 800 km, and a period of 120 minutes, which will provide communication window of up to 13 minutes. The average contact duration is estimated to be 10 minutes. PI-SAT is required to have ability to send and receive 50 Kbytes of housekeeping data and sending picture image of 800 Kbytes within this duration time. This translates to a data rate of 400 bps for housekeeping and 1.2 kbps for image data handling. In order to obtain image data with good quality, goal is set to have maximum bit error rate of 1.0e-5. 800 Kbytes of image data requires the transmitter power to be relatively high in order to get enough power to the receiver back on Earth. The limited electrical power onboard limits the transmitter effect and the overall power consumption of the radio. We arbitrarily choose 4W of radiated power, which at a typical amplifier efficiency of about 40 % translates to about 10 W DC power when transmitting image data. The transmitter will only be used in periods of maximum 13 minutes and can be power switched to beacon mode, which will require only 0.5 W.

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13.2.1 Predicted performance / Link budget Figure 13.1 presents link budgets of PI-SAT command and telemetry, which considers design requirements. It contains frequency, gain of antenna, transmission loss, Eb/No and other important parameters, which represent performance of telemetry link. Link Budgets for PI-SAT

ITEM SYMBOL UNITS UPLINK DOWNLINK beacon(housekeeping) FREQUENCY f GHz 2.30 2.40 2.32 TRANSMIT POWER P W 50.00 4.00 0.50

GAIN OF ANTENNA G dbi 14.00 4.00 4.00 DATA SIZE D kbyte 50.00 800.00 ~ DURATION OF CONTACT t minute 10.00 10.00 ~ DATA RATE (HOUSE KEEPING) Rh bps 400.00 ~ 400.00 DATA RATE (PICTURE) Rp bps ~ 12000.00 ~

ANTENNA EFFICIENCY h % 60.00 40.00 40.00

max path length km 3000.00 3000.00 3000.00 LOSSES

TRANSMITTER LINE LOSS Ll dB -0.50 -1.00 -1.00 SPACE LOSS Ls dB -171.72 -169.59 -169.26 SYSTEM NOISE TEMP k K 650.00 135.00 135.00 polarization loss La dB -0.30 -0.30 -0.30 OUTPUT

EIRP ~ dBW 30.49 9.02 -0.01 E0/N0 ~ dB 36.92 19.64 25.71 C/N0 ~ dB-HZ 62.94 60.43 51.73 BIT ERROR RATE BER ~ 1.00E-08 1.00E-06 1.00E-08 Desired BER BER 1.00E-07 1.00E-05 1.00E-07 required Eb/No dB 15 10 15

implementation loss dB -2 -2 -2

Margin dB 19.92 7.64 8.71 Figure 13.1. Communications budget.

Eb/No is the measure of signal to noise ratio for a digital communication system. It is measured at the input to the receiver and is used as the basic measure of how strong the signal is. In our link budget Eb/No has a margin of 7 db to 20 db, which means its input is strong enough to satisfy required maximum bit error rate. The reason of choosing the frequency will be discussed below.

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13.2.2 Frequency Selection There are many factors to consider when choosing a frequency for the satellite. If we are using omni directional antenna, the ideal frequency is as high as possible. Also the antenna on the satellite will be easier to construct, smaller and more efficient at higher frequencies. Another factor is the frequency allocation – what frequencies that are available. The amateur frequency bands will be far easier to get permission for. AMSAT is in charge of worldwide coordination of the frequency bands around 145, 435, 1215 and 2400 MHz. Even though it is possible to buy economical lower band, 2400 MHz band were chosen for PI-SAT considering the advantage of using high frequency band. Currently Purdue ground station doesn’t have ability to communicate in this band range, so purchasing of new telecommunication equipment will need to be considered. 13.3 How communications supports concept of operations Previous section shows that predicted communication performance satisfies the requirements, which is to have a bit error rate less than 1.0e-5. This performance is good enough to receive picture of required quality. The flight plan of PI-SAT will be discussed in this section. Since we are using omni directional antenna, the satellite and ground station can start contact as soon as the satellite appear above horizon (~3 degree elevation). The footprint is calculated to be 2.9e7 km2. When the satellite is in outside of contact area, the telemetry system is in standby mode, which doesn’t consume power. It will switch to beacon mode right before AOS time (the time when the satellite rises above horizon). Beacon mode contains its housekeeping data. The frequency is 2.32 GHz, and consumes 0.5 W of power. As soon as ground station receives the beacon signal, it sends satellite a command to switch to download mode. This signal is sent with frequency of 2.30 GHz, power of 50W. After the satellite switches to download mode, it starts sending image picture to the ground station with frequency of 2.4 GHz, power of 4 W. After finishing the download, ground station sends satellite next flight plan and command to switch to standby mode. This is the process taken to accomplish our mission. Figure 13.2

Figure 13.2 shows the solution to deal with flipping of the satellite. We have to consider the worst case, which satellite flips completely upside down. The satellite still needs to be able to communicate with the ground station in order to fix this problem. Adding a simple mechanism to our satellite can solve this problem. The solution is to rotate the antennas 90 degrees. Rotating the antennas allows for contact if the satellite is pointing away from the Earth.

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13.4 Subsystem components The satellite communication subsystem provides the transmission, reception, and conditioning of housekeeping data and mission data signals for the PI-SAT. It consists of 3 major component groups that provide a variety of functions: • S-band omni directional Antenna • S-band Receiver • S-band Transmitter 13.4.1 S-band Antenna

Figure 13.3. Radiation characteristics of antenna.

From the restriction of attitude control section, the antenna of PI-SAT is required to be high gain omni-directional antenna. ANT-4-R, the product of DTC corporation is a 4 dbi omni directional RHCP Rod antenna with 0° elevation. This antenna is right hand circularly polarized antenna for 1.7-1.85GHz and 2.4-2.6 GHz. This is used for both transmit and receive applications. The radiation characteristics are shown in Figure 13.2.

2060 MHz azimuth 2060 MHz, elevation

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It provides 0 degree angle circular coverage in the horizontal plane when mounted in the vertical position. In order to obtain wider coverage, PI-SAT will have two receiving antenna and two transmission antenna on the face facing the Earth. Specification are shown in Figure 13.3.

Type Terminated Bifilar Helix

Polarization Right Hand Circular

Gain 4 dBi

Horizontal Pattern Omnidirectional

Vertical Elevation 0°

Beamwidth: Horiz Omni Vert: 50°

MHz

MHz

Feed Impedance 50 Ohms,

TNC Male on bottom of rod

Required No

13.5 X 1.75" (Low),

11.5 X 1.75" (High)

Weight 5.2 oz

Specifications

Bandwidth

Size

Figure 13.4. Specs for antenna.

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13.4.2 Receiver

Figure 13.5. Receiver.

13.4.3 Transmitter Figure 13.6. Transmitter

Receiver used for PI-SAT is expected to fulfill following requirements.

• Compact and light weight • Minimal power requirements • Compatible with antenna

OmniSTAR 3100LM receiver module is product of OmniSTAR OEM technology, which satisfies above requirements. The detailed specifications are shown below, and the receiver is shown in Figure 13.4.

Environments Operating temperature: -20~60 degree Non operating: -40~85 degree Humidity: 95% non condencing Shock: MAX 4G, 5-20mSec Power Power supply: 10Vdc to 22Vdc Power consumption: 350mA at 12Vdc

Physical characteristics Weight 650g Size: 230*110*35 mm(L*W*H) Control: Power switch, command port

Transmitter used in PI-SAT has following requirements

• Compact and light weight • Minimal power requirements • Compatible with antenna • High data rate

The SSTL S-band transmitter module is product of SURREY, which satisfies above requirements. The specification is shown below.

Power requirements: 5V, 680mAEnvironmental: -20 to 50 degree Weight: 500g

Frequency 2.2-2.45 GHz Dimension: 160-120-20 mm Data rates: 9.6Kbps-20Mbps

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13.5 Trade studies / Comparison of components Antenna Several different omni directional antennas were considered before it was fixed to particular one. Figure 13.6 shows Starlink Invicta 210S, L-band omni directional antenna of OmniSTAR corporation. The dimension is smaller than ANT-4-R, but the gain of the antenna was not big enough to transmit image data (1 dBi). Figure 13.7 is ANT-2, a S-band omni directional antenna of DTC. The range of frequency is 2.4-2.484 GHz, which fits to our frequency selection. The weight is 1/5 of ANT-4-R Even though it having many advantages than ANT-4-R, the gain was less than it is required (2 dBi). Figure 13.8 is high gain omni directional antenna of Winncom Technology Corp. It has 13.5 dBi of gain, which is extremely higher than other antennas. The disadvantage of using this antenna is its big dimensions. It has 0.7 m of high, which is bigger than satellite itself.

Figure 13.7. Starlink Invicta.

Figure 13.8. ANT-2.

Figure 13.9. Winncom Antenna.

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Receiver Figure 13.9 shows S-band Receiver of Summation Research Inc. Frequency range fits to our requirement, and the size is slightly bigger than OmniSTAR 3100LM. The only disadvantage is its data rates. The data rate is 1.2 Kbps to 1.6 Mbps, which is too high for our housekeeping data rate (400bps).

Figure 13.10

Transmitter Figure 13.10 shows SAT HDR GOES transmitter module is product of SDI. Compared to SSTL, it consumes more power, and weight more. The advantage of this transmitter is that it is more resistible to harsh environment. The biggest disadvantage is that its maximum data rate is 9600bps, which is not enough to send picture image in within 10 min.

Figrure 13.11.

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Command and Data Handling 14.1 Requirements The command and data handling system (CDH) is responsible for the communication and coordination of all the connected subsystems of PI-SAT. It commands the onboard computer. The system is used for processing and distributing commands onboard the satellite. It should be able to respond to requests from the ground station obtained through the communication subsystem, and should be able to service the needs of each subsystem. Except for uploaded flight plans from the ground station, this system should make the satellite totally autonomous. Figure 14.1 lists the requirements of this system.

RequirementsInterpret flight planStore Data -Operating System ~ 1.2 MB -Image File ~ 765 kB -Housekeeping Log ~ 1-10 kBControl power modes of subsytems and their compnentsControl ADCS to orient itself autonomouslyTransfer data from camera to memory and memory to COMTake temperature measurements for logCheck amount of powerControl COM to send data to ground station

Must be able to…

Figure 14.1. Requirements for CDH.

14.2 Subsystem Components CDH is divided into 2 sections: Hardware and Software. The hardware component is comprised of the computer, the different types of memory, framegrabber for camera, and interfaces between all the different subsystems. The software is just comprised of the operating system and the flight plan for the satellite. Figure 14.2 is a functional diagram of how CDH is connected to the different subsystems of the satellite.

Figure 14.2. Flowchart of the control layout.

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• Command and Data Handling (CDH) – Controls the onboard computer by receiving, processing and distributing data to and from the other subsystems.

• Power (PU) – Charges batteries, checks power levels, and contains power conditioner.

• Batteries (Batt) • Attitude Control System (ACS) – Controls the orientation of the satellite using

sensors and magnatorquers. • Sun Sensor (Sensor) • Magnatometer (Actuator) • Communications (COM) – Communicates back and forth from ground station. • Receiver/ Transmitter (Rx/Tx) • Payload (CAM) – The camera.

14.2.1 Hardware The heart of the satellite is the onboard computer. Since the satellite is not very complex, the computer will not have to be too powerful. Figure 14.3 shows RTSCM, the chosen motherboard for PI-SAT, made by Maxwell Technologies Microelectronics.

Figure 14.3. Manufactured Maxwell Technologies Microelectronics.

This radiation-hardened module has a 20 MHz processor, and has 8 MB of onboard SRAM. It requires less tan 5 watts at full power, and 1.4 for low. At a weight of 1.0 kg, it can operate between -35o to 75o C. This module can sustain a radiation dose up to 60 Krad. The complete specs for this computer are in Appendix 14.2a. The camera for the satellite is a CMOS camera, so somehow the system must be able to “grab” a picture from the video being sent from the camera. Figure 14.4 shows EHDvideoPC-C, the framegrabber that was chosen, made by EHD.

Figure 14.4. Manufactured by EHD.

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The framegrabber is a PCMCIA card with two composite and one S-Video inputs. It requires 5V ± 5%, with an operating current of 250 mA. The specs for the framegrabber are in Appendix 14.2b. This framegrabber, which only weighs 35 gm, cannot be directly connected to the chosen motherboard, so an adapter is needed. Figure 14.5 shows MiniModule PCC, an adapter that can be used to connect the framegrabber to the motherboard.

Figure 14.5. Manufactured by Ampro Corp.

The adapter is capable of holding two PCMCIA cards, if needed. It only weighs 102 gm, and uses 45 mA at 5V ± 5%. This adapter can be connected to the motherboard. The complete specs for the adapter can be found in Appendix 14.2c. 14.3 Impact of Space Environment At an altitude of 800 km, PI-SAT should experience approximately 1 Krad of radiation per year. Since, the computer that was finally chosen for the satellite is radiation hardened, it is unlikely, but not impossible, for the satellite to malfunction due to radiation effects. A trade study was done, which is explained in Section 14.5, about the use of a different computer that would have required a very thick and heavy aluminum box for protection. 14.4 Operations and Functions The CDH system is responsible for most of the decision making on the satellite. When a flight plan is uploaded and stored in the computer, it will read the plan to determine when and what it is supposed to do. There are several commands that could be written into the flight plan. Those could be: -Check power status -Check temperature status -Take picture -Initialize ACS -Transmit data Figure 14.6 shows the process the CDH system goes through with the flight plan.

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Figure 14.6. Reading the flight plan.

When the command to take a picture comes up in the flight plan, a few little jobs must be run before and after the picture is to be taken. Figure 14.7 shows the process the system will go through in order to take a picture.

Figure 14.7. Taking a picture.

One important step in this procedure involves checking the power. If there isn’t enough power to handle the task of taking a picture, then a picture shouldn’t be taken. Therefore, if there wasn’t enough power, the system would log the low power and that it had to end the task. When the satellite initially gets to orbit, it will need to turn itself on and boot up the software. Figure 14.8 shows the process of booting up the satellite.

Figure 14.8. Initializing the satellite.

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The system is also responsible for sending the data down to the ground station. When this command comes up in the flight plan, the computer is responsible for sending the data from storage to the transmitter. 14.5 Trade Studies There are two possibilities that were analyzed for the command and data handling system. Both systems were very similar, except for the computer. The first system used CoreModule 3Sxi, as shown in Figure 14.9, as the main computer for the satellite.

Figure 14.9. Manufactured by Ampro Corp.

The module has a 25 MHz 3786SX-compatible processor, and can contain up to 8

MB of onboard DRAM. It requires 5V ± 5%, 500 mA when active and 160 mA when asleep. At a weight of 85 gm, it can operate for 0o to 70o C. The complete specs for this computer are in Appendix 14.5. The problem with this computer is that it isn’t radiation hardened, and it would have required an aluminum box approximately ¼” thick, similar to the one shown in Figure 14.10.

Figure 14.10. Aluminum box to shield from radiation.

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This is just unacceptable. One of the requirements for the satellite is to try and

keep the mass as low as possible, and with this box it would be very difficult to do. Figure 14.11 shows a comparison of this system, System A, and the chosen system, System B.

Mass 7.85 kg 2.0 kgPower 4 W 6.5 W

Advantages Smaller in size Can withstand radiationFaster Processor Larger operating

temperature range.

Disadvantages Would probably need redundant Much more expensive. system in case of radiation malfunction.

System A System B

Figure 14.11. Comparison of the two systems.

After comparing the two systems, the less weight and capability to withstand radiation made System B the better choice. Maybe down the line, cost will be a larger factor and cause System A to be the better choice. So, for now, System A is shown here as another possibility for the command and data handling system.

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15 Power 15.1 Major Requirements The major requirements of the power system are the communications, attitude determination and control, command and data handling, and payload subsystems. The satellite’s power system must be able to supply sufficient power to the satellite subsystem to successfully perform the mission. The power system must also be able to maintain the lithium-ion cells for the duration of the mission. 15.2 Operating Modes

There will be three major operating modes for the satellite, which need to be considered from a power standpoint. The mode that the satellite will be in the most is the station-keeping mode. This mode will be used when the satellite is waiting to receive data and maintaining it position. During this mode the satellite will be drawing it lowest amount of power, will allow the satellite to recharge it batteries. The second mode will be the data transfer mode. This mode will be used to send the data from the satellite to the ground station. During this mode the pointing of the satellite will have to be maintained along with the communications link for the time of the pass over. The finally mode is the full power mode or picture taking mode. As the name implies this mode is used for the taking of the pictures, which is the main part the satellite’s mission. During this mode the satellite will need to be taking a picture and maybe even communicate with the ground station at the same time. This mode will make up the smallest part of the satellite’s mission. 15.3 Subsystem Components The power system is made up of three major parts the power generating solar cells, the power distribution system and the batteries as the power storage system. The solar cells will provide the power for the satellite. Four Gallium arsenide (GaAs) solar cells will provide the satellite with a minimum of 70 watts of power in sunlight and a maximum of 100 watts. This power will then be transferred to the power distribution system. Which will route the power to the 8 lithium-ion cells, used for power storage or to the satellite subsystems. The power can also be routed from the li-ion cells to the subsystems while in eclipse. The power distribution system consists of a battery charge regulator (BCR) a power routing switch and a power conditioner. The BCR will monitor the charge of the lithium-ion cells and route power to or away from the cells as needed. The power conditioner in figure 5.1 will be used to condition the satellite’s base power of 28 Volts to the voltage required for each subsystem.

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Figure 15.1. Power conditioner for satellite power subsystem.

15.4 Power Budget The power budget was based on the numbers calculated for the other subsystems. The numbers for the communication and command and data handling subsystems have been approximated as closely to exact as possible, due to the lack of exact numbers on those components. After a subtotal was calculated a margin of 25 percent of that subtotal was added to it to get a total power needed for the system. This margin has been added to allow for cable losses and the inefficiency of the power conditioner.

Subsystem Full Power [Watts]

Data Transfer Mode

Station Keeping Mode

Communications 10 10 5 A 3 Mag Torquers [total] 0.54 0.54 0.54 E 2 Sun Sensors [total] 0.1 0.1 0.1 E Magnetometer 0.56 0.56 0.56 E Payload 6 0 0 E CD&H 10 5 5 A Subtotal 27.2 16.2 11.2 Margin [25% of Subtotal] 6.8 4.05 2.8 Total Power needed 34 20.25 14 E=Exact A=Approximate

Figure 15.2. The power budget of each subsystem.

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Power Distibution in Station-Keeping Mode

3 Magnetic Torquers [total]3.86%

2 Sun Sensors [total]0.71%

Magnetometer4.00%

Communications38.82%

Margin [25% of Subtotal]20.00%

CD&H35.71%

Figure 15.3. A pie chart of the power distribution for the station-keeping mode.

15.5 Trade studies / comparisons of components For the power components of the satellite an effort was made to make them as efficient as possible, when considering weight and size and so much the cost. For the battery cells the early estimates of the satellite power were much higher than the current needs. So to provide this power needed and maintain a strict conformation of weight and size Lithium-ion batteries were chosen. We decided to use 8 cells at 3.6 volts each to provide a voltage of 28 volts for the systems voltage, which was needed for our power conditioner and was the norm for smaller satellite power systems.. Along with the batteries, solar cells were needed to provide power for the satellite subsystems along with power to charge the lithium-ion cells. Gallium arsenide was chosen due to its high efficiency. We need to have a smaller solar panel to allow for a non-deployable solar array system. By making it not deployable we have decreased the weight and complexity of the structure while maintain a smaller volume. Finally the power conditioner was chosen for its small size and weight.

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17 Structures and Mechanisms 17.1 Major Requirements The main structure of the spacecraft was designed to be simple to build and relatively inexpensive. There are no exotic materials used in the construction, and it is designed to be built by students. The spacecraft structure is designed to support a total weight of 51 kg through launch and orbit operations. 17.2 Structures Support of Operations The structure of the spacecraft is designed to contain all of the components, which make up the spacecraft, and protect them from the harsh environment of space. The structure must be able to withstand the launch vehicle forces, and protect the components throughout the mission. The structure is most definitely a mission critical subsystem that must function properly for a successful mission. The gravity gradient boom is the only moving part on the spacecraft, and it will be deployed once we are free of the launch vehicle. The gravity gradient boom is manufactured by Surrey Satellite Technology Ltd. and it contains all of the moving parts and mechanisms inside it for deployment. It will be attached to the structure inside the tube shown in Figure 17.1.

Figure 17.1. Top view of satellite.

17.3 Major Loading Conditions The launch vehicle will impart the most force on the spacecraft throughout the mission. Since the spacecraft is designed to withstand the launch loads, it will be able to withstand any of the other smaller loads that will come about through normal mission operations. The launch vehicle is explained in detail in Section 9, please reference this section for more detailed information. The components of the spacecraft will apply significant loads when under acceleration from the launch vehicle, the sheets and beams had to be iterated through the use of ANSYS to find an optimal balance between weight and strength, this is discussed in more detail in sections 17.4 and 17.5.

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17.4 Spacecraft Structure Concept and Material Selection 17.4.1 Support Beams The main support beams are Aluminum 6061 hollow square tubes with the dimensions given in figure 17.2.

Fig. 17.2. Cross section of Aluminum support beam (cm).

Fig. 17.3. Cross section of Steel support beam (cm).

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17.4.2 Stringers The stringers that are used throughout the spacecraft are Aluminum 6061 and are described in Figure 17.4.

Fig. 17.4. Cross section of stringers (in).

17.4.3 Sheets The sheets are Aluminum 6061, and are 0.16 cm thick. 17.4.4 Material Properties The material properties of Aluminum 6061 are given in Figure 17.5.

Material Properties of Aluminum 6061 E 69x109 Pa

Density 2800 Kg/m3 ν 0.33

Fig. 17.5. Table of Al 6061 properties.

The material properties of Steel 1005 AISI are given in Figure 17.6. These are the beams that are on the payload adapter face.

Material Properties of Steel 1005 AISI E 200x109 Pa

Density 7872 Kg/m3 ν 0.33

Fig. 17.6. Table of Al 6061 properties.

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17.4.5 Structure Layout ANSYS was used to model the structure, and Figure 17.7 illustrates the main support of the structure. Red beams are Steel square beams as in Figure 17.2, green beams are Aluminum square beams as in Figure 17.3, and blue beams are Aluminum L stringers as in figure 17.4.

Fig. 17.7. ANSYS model of beams.

17.5 Predicted Performance The spacecraft was extensively modeled in ANSYS. This provided us with a very detailed analysis of the spacecraft and its behavior. The natural frequency of the spacecraft was determined and the first ten modes are given in Figure 17.8. Stress analysis was also performed in ANSYS, which provided information on weak areas, and also areas where material could be removed, reference Figure 17.9. This greatly assisted us with developing our structure, and many iterations were performed to find an optimal design that met all of the mission requirements. The components were modeled as a point mass that was distributed over a certain area of the sheet. This is illustrated in Figure 17.10, which shows each of the components as a “web” which simulates real components much better than lumped masses.

Natural Frequencies (Hz)

1 91.54 2 101.11 3 131.19 4 132.16 5 138.16 6 142.29 7 148.91 8 160.89 9 164.31

10 196.90 Figure 17.8

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Figure 17.9. Stress at first natural frequency.

Figure 17.10. Component Modeling.

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17.5.1 Safety Factor The acceleration of the launch vehicle is given in figure 9.5, and the maximum

stress is given in figure 17.11. The maximum stress of 2.04e6 Pa occurs at the aluminum sheet, which yields at 1.03e8 Pa. Therefore there is a safety factor of 50.49 during the launch phase. The main driving factor of the spacecraft was the first natural frequency; the stress at the first frequency was 1.78e10 Pa. The high stress occurs at the aluminum sheet, which has an ultimate stress of 2.28e8 Pa, therefore the structure would eventually fail if it were vibrated at its natural frequency of 91.54 Hz. The structure is over designed for the launch loads, but its stiffness is necessary for the natural frequency.

Figure 17.11 – Launch Stress

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17.6 Trade Studies Aluminum was chosen as the main structural material early in the design process because of its high strength to weight ratio. Many other spacecraft performing similar missions also used aluminum, so it is a well-understood and predictable material. Composite materials were considered for some time, however their high cost and low level of experience made them an unnecessary risk for this spacecraft. The face on the spacecraft that attaches to the launch vehicle experiences extreme forces compared to the rest of the structure, so Steel 1005 AISI beams were used on this face. The beams have the same outer dimensions as the other structural beams, but they are thicker and have a higher moment of inertia, reference Figure 17.3.

Through the use of ANSYS the beams, stringers, and sheets were changed to produce the desired results. The shapes of all the structural elements stayed relatively the same, but their size and thickness were varied. The main driving factor was the first natural frequency; it had to be higher than 90 Hz for the Ariane 5 ASAP program, reference section 9 for more information. The stress during launch was not as large as during the first natural frequency, therefore structural failure is not likely. The requirements of the spacecraft made it necessary for the gravity gradient boom to be made by an established and experienced company. We were not intending to design a gravity gradient boom for this mission. Many other similar satellites have used booms made by other companies.

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18 Spacecraft Summary PI-SAT is still in the preliminary design stage, however many of the crucial elements of this student designed micro satellite are identified. The current design meets all of the requirements, and is capable of continuing on to the next stage of development. 18.1 Requirements Compliance

Group Requirement Solution Section Capable of imaging entire earth Sun synchronous orbit 8 Orbit Large amount of time in Sun for

Solar Panel charging Sun synchronous orbit

Launch Total mass must be less than 120 Kg.

Total Mass is 50.5 Kg. 9

Payload Take a Picture of the Earth at a resolution of at least 1 km

Two CMOS cameras and support

11

Achieve and maintain nadir pointing

Gravity gradient stabilization, magnetic control, and nutation

damper

12

Pointing determination to less than 10 degrees

Sun sensors and magnetometer 12

ADCS

Attitude failure recovery Magnetorquer powerful enough to counter gravity

torque

12

COM Send picture image of 800 kb within 10 min. The bite error

should be less than 1%

Utilize 4 dbi omni directional antenna with power of 10 dcW

13

Must have enough memory for operation

Equipped with 8 Mb of RAM 14 CD&H

Computer must be safe from radiation effects

Computer is radiation hardened 14

Power Supply all subsystems with necessary power

Li-ion Batteries, Solar cells and Power Conditioner

15

Thermal Maintain internal temperature of subsystems

Thermal Blankets, and sun synchronous orbit

16

Natural Frequency must be greater than 90 Hz

The natural frequency is 91.54 Hz

17

Spacecraft must adapt to Ariane 5 ASAP

ASAP adapter was designed for spacecraft

17

Structure

Spacecraft must withstand launch loads

Safety Factor of 50 17

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18.2 Cost Estimation

Since the satellite is still in development, a good cost estimation is difficult to complete. However, the launch price is estimated to be $10,000 per pound, therefore the launch cost alone is approximately 1.2 million dollars. All of the hardware is currently available from other companies, and can be purchased for a reasonable amount of money. The structure is simple and could most likely be manufactured by students at Purdue. With these constraints, we estimate the currently designed satellite to cost less than 2 million dollars to build and launch, which is a reasonable price for a micro satellite. 18.3 Further Development These are the areas that would require significant improvement if an operational satellite was to be developed:

•Software development •Analyze radiation effects of hardware more •Thermal •Analyze effectiveness of thermal blankets •Attitude control •Develop nutation dampener •Structure •Decrease mass and size

19 References (To be included) 20 Appendix (To be included)