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In recent decades, the number of satellites being built and launched into Earth’s orbit has grown immensely, alongside the field of space engineering itself. This book offers an in-depth guide to engineers and professionals seeking to understand the technologies behind Low Earth Orbit satellites.

With access to special spreadsheets that provide the key equations and relationships needed for mastering spacecraft design, this book gives the growing crop of space engineers and professionals the tools and resources they need to prepare their own LEO satellite designs, which is especially useful for designers of small satellites such as those launched by universities. Each chapter breaks down the various mathematics and principles underlying current spacecraft software and hardware designs.

### Chapter 1. The Space Environment

Space environment and related matters as they apply to spacecraft hardware design will be discussed in this chapter.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 2. Satellite Missions

The most frequent orbits are shown in Fig. 2.1. Polar orbits are those where the plane of the orbit passes through the poles. They have inclinations of 90° and are usually circular. Because the Earth rotates under the orbit, these satellites can survey the entire Earth.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 3. Orbits and Spacecraft-Related Geometry

Acceleration of gravity, g(H), at an altitude, H, varies inversely with the square of H. The equation for g(H) is given below. Its value at the surface of the Earth is go and is approximately 32.2 ft./sec2 or 9.8 m/sec2. R is the radius of the Earth and H is the orbit altitude, both in km.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 4. Electric Power Subsystem Design

The Electric Power Subsystem (EPS) of a satellite is a heavy and expensive subsystem. It is often about 25% of the weight and 25% of the cost of a spacecraft. Electric Power is also often underestimated, resulting in insufficient power to support the “mission creep” requirements of the spacecraft. The EPS design procedure is outlined below. Each of the steps will be discussed and illustrated in more detail later.
(A)
Determine the Required Spacecraft Orbit Average Power
• List all of the electronic components of the satellite and the voltages and currents that each component requires
• Determine the power drawn by each component in each of the spacecraft operating modes. Augment these by the appropriate DC/DC conversion efficiency to obtain the OAP drawn from each voltage source in each spacecraft operating mode.
• Determine the peak OAP required

(B)
Determine the Battery Capacity Required and Choose the Battery Bus Voltage
• Based on the power drawn during the eclipse (and the maximum eclipse duration), determine the battery WH requirements
• Select the battery cells that will be used
• Applying the battery output vs. input efficiency, determine the battery WH used during the eclipse
• Select the maximum Depth of Discharge below which the batteries should not be discharged. Apply this, and a large safety factor, to obtain the battery WH to be installed.
• Choose a battery bus voltage and divide by the cell voltage to determine the number of cells in series (in a string) of cells. Divide the total battery current by the current each parallel string will supply to determine the number of parallel battery strings.

(C)
Select a Solar Panel Configuration and Compute the OAP it can Supply
• Select the solar panel configuration (the orientations and areas of each panel relative to the spacecraft axes). Also, determine how each panel will be stowed and released.
• Compute the instantaneous power generated by each panel as the spacecraft moves around an orbit. The total power vs. time is then computed, as is the OAP.
• Repeat this for all Beta angles (the angle between the sun line and the orbit plane) to determine what the minimum OAP is. Ensure that the minimum OAP generated is equal to or greater than the spacecraft OAP required.

(D)
Draw the EPS Block Diagram
• Given the panel configuration, the various required voltages and the number of battery strings and cells per string, the EPS block diagram can now be drawn.
• Consider which groups of components should be turned ON/OFF on command, and whether the switch to turn these groups ON/OFF should be ahead or after the respective DC/DC converters that supply the voltage to the group.

(E)
Miscellaneous EPS Design Steps
• Often, an EPS computer is included to collect telemetry regarding the state of health of the EPS, the battery capacity status, component temperatures and EPS status. This computer may also be used to turn ON/OFF power to the various electronic components.
• The Separation Switch that signals release from the launch vehicle and the start of spacecraft operations is also part of the EPS. The functions enabled or disabled to ensure that no electric power is drained from the spacecraft prior to launch are used to determine where in the spacecraft circuit the Separation Switch should be located.

George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 5. Spacecraft Communications

RF communication is used to command the spacecraft from the ground, to have the spacecraft send Health and Status (Telemetry) about the condition of the spacecraft and to send payload data to the ground station(s). This chapter discusses spacecraft frequencies, communication link margins, Bit error rates and RF hardware.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 6. Spacecraft Digital Hardware

Digital computer selection for spacecraft applications is not as big of a problem as it used to be, because computers have increased in speed and in the size of their memories. Thus, these are no longer major factors in the selection process. However, radiation hardness, sensitivity to SEU, power consumption and I/O capabilities are still properties that influence the choice of digital processors. A list of factors that should be taken into consideration is given below:
• Architecture (one central or a distributed set of computers)
• Software Operating System
• Instruction Set
• Number of Bits Per Word
• I/O Capability
• Reliability
• Redundancy
• Memory Size
• Speed
• Cost
• Weight
• Power Consumption
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 7. Attitude Determination and Control System (ADACS)

The ADACS is one of the most expensive subsystems of a spacecraft. Over-specifying ADACS performance requirements can easily “bust the budget.” A common mistake is to call for more stringent accuracy than is required by the payload. For example, if the mission is to simply scan the atmospheric spectrum from the Northern Hemisphere, then pointing much more than 5° is hardly required. Specifying pointing accuracy 1 or 2 orders of magnitude better rapidly leads to expensive equipment, such as star trackers, when a simple momentum-biased system with Earth sensors would suffice.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 8. Spacecraft Software

Spacecraft software is discussed here in terms of (1) the functions the software has to perform, (2) the software architecture and (3) the manner in which the functions are performed. While the functions the software has to perform are generic, there can be different architectures and different methods of implementing the functions. Here, a point design is described, a design with a lot of flight heritage over many different spacecraft. The spacecraft missions in this example are:
• Store and Forward communications
• Collecting data from and commanding unattended remote sensors
• Imaging designated targets and downlinking image data from the spacecraft to ground stations
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 9. Spacecraft Structure

Selection of the most suitable spacecraft structure depends on many factors. The main ones are (1) the launch vehicle payload envelope and interface separation system dimensions, (2) the launch vehicle loads to the spacecraft structure, (3) the weights of the spacecraft bus and payload components, (4) thermal design and ability to get rid of incident and internally generated heat, (5) the way solar arrays will be mounted, deployed or rotated to provide the required Orbit Average Power, (6) instrument pointing and (7) any other special requirements, such as the possible requirement to keep propulsion, bus and payload separated from one another.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 10. Deployment Mechanisms

Various parts of the spacecraft often have to be deployed on orbit. These may include solar panels, antennas, gravity gradient booms and various other booms and instruments of the payload. There are two main rules for designing deployables.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 11. Propulsion

In LEO spacecraft, the main uses of a propulsion system are to raise or lower the orbit and to maintain the spacecraft on station in a constellation of spacecraft. The Rocket Equation, given below, describes the fundamental property of a propulsion system. It gives the magnitude of change in spacecraft velocity as a function of the spacecraft and expanded fuel masses.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 12. Thermal Design

The altitude, inclination, epoch and angle of the orbit plane to the Sun (the Beta angle) determine the thermal environment of a spacecraft. In the direction of the Sun, the heat incident on a spacecraft can be huge, while on the side opposite the Sun, the spacecraft faces cold space. The objective of the thermal design is to bring these temperatures into a range where the spacecraft components can safely operate, and to ensure that internally generated heat from the spacecraft components is conducted or radiated out to maintain all spacecraft temperatures within the operating temperature ranges of its components. Typical component temperature operating ranges are given in Fig. 12.1. The goal is to design for component temperatures in the 10°–20 °C range.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 13. Radiation Hardening, Reliability and Redundancy

There are at least three factors that affect spacecraft orbit life. These are (1) the amount of radiation protection provided for the orbit altitude of the spacecraft, (2) the reliability of the components and of the spacecraft system and (3) the redundancy built into the spacecraft.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 14. Integration and Test

Integration is the process of testing the components of the spacecraft, assembling them, putting the completed spacecraft through functional testing, then performing thermal, vibration and thermal vacuum tests. After that, the spacecraft is ready to be shipped to the launch site, where it is functionally retested to ensure that no damage was done to the spacecraft. Then, the spacecraft is integrated with the launch vehicle.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 15. Launch Vehicles and Payload Interfaces

There are two essential ingredients one must have in interfacing with launch vehicles: money and patience. The selection and manifesting process is long, schedules are uncertain, and a launch failure on an earlier flight could introduce additional years of delay.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 16. Ground Stations and Ground Support Equipment

Ground stations (1) command the spacecraft(s), (2) collect, display and analyze spacecraft telemetry about its state of health and (3) command payloads and retrieve data from them.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 17. Spacecraft Operations

Here, the methods of operating a spacecraft and its payload to (1) schedule events, (2) display health and status telemetry, (3) command changes in the setup of spacecraft or payload parameters, (4) operate the propulsion system, if any and (5) resolve and correct spacecraft operating anomalies will be discussed.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 18. Low Cost Design and Development

In 1995, I wrote a chapter on the RADCAL spacecraft for Reducing Space Mission Cost, a book in the Space Technology Series. The book was edited by James Wertz and Wiley Larson. In the chapter, I wrote a section on the practices used to come up with a very capable and complex spacecraft at a very low cost. The schedule (from contract award to launch) took one day less than one year. In reviewing that chapter now, after having developed 34 spacecraft, I must say that each procedure for keeping costs down described in that chapter is true today. For this reason, this chapter is largely taken from that 1995 book. It is just as true today as it was then.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 19. Systems Engineering and Program Management

The Program Manager has the responsibility to deliver a spacecraft that meets customer functional requirements and specifications, and is constructed within available funds and within a specified schedule. He or she also ensures that the spacecraft meet launch vehicle and ground station interface requirements, and that the spacecraft, when on orbit, can perform the specified mission.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra

### Chapter 20. A Spacecraft Design Example

This example of a spacecraft design problem is given so that the reader can put to work what he or she has learned by reading this book. While the “Requirements” are clear, they are not very specific. There is not one good answer or design to meet the requirements.
George Sebestyen, Steve Fujikawa, Nicholas Galassi, Alex Chuchra