Since 1963 and the launch of Syncom II, the first geosynchronous Earth orbit (GEO) telecommunications satellite, the GEO orbit has become significantly more crowded (“Harold Rosen,” 2000). With over hundreds of satellites now crowding the orbit, multi-national satellite operations agencies are researching new methods of tracking GEO satellites and debris (“Traffic management in,” 2008). Even with several ground-based space tracking systems across the world, many items in the GEO belt are lost or untracked. The world needs a better system of tracking objects in GEO. Military and civilian agencies in the United States have identified a requirement for a “more precise means of tracking objects in the GEO belt.” The intent of this paper is to detail the engineering requirements associated with a space-based GEO tracking capability, the Supersynchronous Tracking Device (STD).
STD will be a system designed to provide mapping the geosynchronous belt in order to provide more detailed understanding of the GEO belt for collision avoidance. This mission is designed to provide Space Situational Awareness (SSA) to a multi-national forum. As a secondary mission, STD will track U.S. high-interest objects in the geosynchronous belt. For these missions, STD is defined as a satellite system that will operate at a super synchronous altitude, have a sensor capable of mapping and tracking objects at GEO, provide a 24/7 relay data to a ground station, and provide on-board filtering of data in order to prioritize change detection.
Stakeholders are identified as any company involved in the ownership or operations of space systems GEO belt and US space tracking agencies. As the need for GEO belt mapping gains dimension as a global necessity, STD will offer a new solution for international integration of space situational awareness (SSA). However, as the need for super synchronous tracking of GEO systems is not yet fully understood, STD will initially offer SSA to U.S. and coalition partners. STD architecture will be comprised of a satellite system, ground system, and a command and control system. The satellite system will map the geosynchronous belt to provide a more robust understanding of the locations of all GEO systems. Furthermore, the satellites within the STD architecture will provide tracking of high-interest objects in the geosynchronous belt to stakeholders involved in owning and operating in the GEO belt and US space tracking agencies. A brief description of the STD initial requirements architecture is found in Appendix 1. Of the two architectures analyzed, STD Simplex II is the option that best suites the stakeholders. Simplex II is comprised of Space, Ground, and Launch segments. The space segment has been divided into subsystem requirements to include: more complex systems – anticipated to be larger than 500 kg, smaller constellation of satellites – anticipated to be less than 10 satellites, and a higher quality of end-product due to a larger available bus and more space for on-board processing. The ground segment has been divided into subsystem requirements to include: minimal command and control supports – fewer dedicated ground antennas are anticipated, and minimal tracking capabilities due to the large size of the satellites and slow motion (with respect to the earth). The launch segment has been divided into subsystem requirements to include: a very specific type of launch vehicle to transport a large payload to super synchronous orbits and utilizing launch vehicles more frequently as each space system will be launched individually. Each one of these segments – space, ground, and launch – will be discussed in detail in this document. All of the requirements for the subsystems, key interfaces, and design choices will be discussed to identify the newfound intricacies of this new type of system. The sophistication of STD is largely attributed to the space segment. As a new endeavor in the domain of space, this segment’s requirements are as unique as the idea as utilizing the super synchronous orbit in an operational capacity. However, many of the payload requirements are drawn from other systems with similar types of sensors. Space segment requirements are as follows:
Design choices of STD are derived from the operational constraints of the sensor. Therefore, the following requisites cover the design choices approved for STD:
STD shall be a spacecraft equipped with a bus designed to support these payload design choices. Therefore, the STD spacecraft will be designed so that it supports at least 2,000 watts to the payload and supporting systems. Due to the low number of satellites in a full constellation of seven, a standard satellite bus designed to last 9-10 years will be sufficient for STD. Solar arrays and electrical subsystem components shall be designed to interface equally with the physical design of the bus as well as the power demands of the satellite. Overall structure shall be a common platform, designed for maximum compatibility.
The Structures and Mechanisms Subsystem (SMS) of STD shall provide the framework to literally hold the spacecraft together. The SMS must adequately support every other subsystem of the spacecraft to ensure successful functioning of the satellite. In addition, the SMS accounts for the launch vehicle adapter, which guarantees the spacecraft can not only attach itself to the launch vehicle but also withstand the physical challenges associated with launch and ascent of the vehicle. All of the various components of a satellite must be mounted to the structure of the spacecraft. The structure shall also serve as a shield to the satellite’s internal equipment from the harsh space environment, including radiation and extreme temperatures. Building the most sensitive components as deeply as possible within the structure further shields the equipment from harmful environmental factors. Lastly, the mechanisms of the spacecraft facilitate the separation and deployment, both mechanically (such as spring-loaded or electronically-driven) and explosively (utilizing pyrotechnics), of the satellite’s deployable parts. The primary structures shall carry and encompass the large loads of the spacecraft, particularly with respect to the launch vehicle. The secondary structures, such as trusses, beams, booms, platforms, and walls make up the chassis for STD. Solar arrays are also considered secondary structures of STD and must be developed to minimize torques and other forces. Lastly are the tertiary structures, the non-weight bearing elements such as wiring harnesses, fuel lines, connector panels, support brackets, and other various connector types and component housing. The SMS for STD must meet several more specific requirements. The SMS shall be constructed of materials which will be able to withstand the applicable pressures and forces of testing, launch, orbit rising, and on-orbit steady-state operations. The SMS shall also ensure the space vehicle operates within established stabilization parameters. The structure shall be able to support a gimbaled sensor, ensuring the sensor has an unobstructed view of its observation targets. The mechanisms shall be able to properly deploy the required equipment necessary for successful operations. The structures and mechanisms of a spacecraft are interlaced throughout the entire vehicle, interfacing with the attitude determination and control subsystem (ADCS), the command and data handling subsystem (C&DH), the telemetry tracking and commanding (TT&C) subsystem, the propulsion subsystem, the power subsystem, and the launch segment, as well as the satellite’s payload (National Aeronautics and Space Administration (NASA, 2003). Particular attention must be paid to the gimbaled sensor onboard the STD payload. Sensor-mounting platforms must be designed and manufactured to satisfy rigidity specifications, and gimballing mechanism must accommodate sensor pointing requirements. The STD satellites will use the standardized Lockheed Martin A2100 geosynchronous telecommunications satellite bus (Lockheed Martin, 2005). The advantages of utilizing a pre-existing frame are a drastically-reduced schedule due to much shorter design and development phases, and potentially lower costs depending on the amount of additional specialized component fabrication required. The design of the A2100 provides a more streamlined and less expensive architecture, weighing less than other comparable models. The A2100 is designed to be a more powerful bus, capable of generating over 15 kW. This increased power is due to the inclusion of more-efficient “pleated shade” solar cells and better heat dispersion. Multiple A2100 configurations exist, such as the A2100A, A2100AX, A2100AXS, etc., depending on mission requirements. The STD will utilize the military-approved A2100M, the same bus currently planned for use in the Navstar, Mobile User Objective System (MUOS), and Space-Based Infrared System (SBIRS) programs. The structural design and materials must meet strength, density, and rigidity standards. They must also be able to withstand the physical aspects of space, such as those caused as a result of extreme temperature changes. Additionally, the materials must be able to resist corrosion and unwanted conductivity. The STD SMSs will be constructed of a combination of homogenous aluminum and magnesium materials, which will better suit the satellites to withstanding launch forces, vibrations, lateral loads, and acoustic noise. Aluminum is a very low-density material which translates to a high strength-to-weight ratio and adapts well to space environmental factors. Magnesium is also a low-density material and provides added stability for stiffness and rigidity requirement satisfaction. Utilizing composite materials will also maintain a suitably stiff platform for the payload’s sensor. A key element to which the SMS contributes is the weight of the spacecraft. The weight of each and every component must be minimized as much as possible. The structure of the satellite itself can account for 20% of the overall weight of the vehicle (Wertz & Larson, 1999, Chapter 11). Similarly, the mechanisms of the spacecraft may be upwards of 10% of the total weight. Weight is an integral factor to the selection of launch vehicle used as well as the amount of propellant required onboard the spacecraft. Weight can be further minimized by designing components as close together as possible, decreasing the amount of wire bundles, wiring harnesses and welds required. Additionally, the SMS must be designed to optimize the dry mass-to-propellant mass ratio (Wertz & Larson, 1999, Chapter 11). There are numerous environments that must be taken into account when designing and developing the structures and mechanisms of the spacecraft: construction, testing and handling, launch, ascent, and on-orbit operations. Constructing the other satellite subsystems simultaneously with the SMS will help to streamline the process as well as potentially helping to meet budget and weight constraints. In addition, the possible need to adjust or replace individual components during testing will be taken into consideration during manufacturing. The launch of a booster creates tremendous forces for its payload, such as constant and intermittent acceleration, acoustic noise, vibrations, and other load factors. All of these elements must be considered during the design phase and mitigated by the SMS. The spacecraft launch vehicle adapter must be designed flawlessly and perform its function correctly in order to preserve the integrity of the satellite within the enclosed fairing throughout launch and ascent and to ensure proper satellite separation from the booster and fairing at the appropriate time. Once the satellites reach their intended orbits, the mechanisms must take over and deploy the necessary items. The solar arrays will be deployed using the Solar Array Drive Mechanism (SADM), which incorporates both explosive, pyrotechnic ordnance release mechanisms as well an electrical drive motor, in order for the most efficient use of energy. A three-axis stabilized structure, although more complex and expensive, provides a stable and maneuverable platform for on-orbit mission operations. The STD SMS will consist of such a platform, incorporating both active and passive stabilization techniques to include thrusters, four reaction wheels, and magnetic torquers.
In order to maintain the position of each satellite in the STD constellation an absolute station-keeping technique will be used. This is where “we maintain each spacecraft within a mathematically defined box moving with the constellation pattern.” (SMAD) Because STD will be at an altitude higher than geostationary orbit the amount of drag from the atmosphere will be minimal as compared to a low earth orbit which would have a higher drag on the satellite. However, each satellite will still be maintained within a defined box. Thrusters and reaction wheels will be used to make sure that the satellite is kept in the box while also maintaining a correct attitude and stabilization. The thrusters will be applied “at the forward edge of the box, the applied delta V is increased, thus increasing the orbit altitude and period and sliding the satellite rearward in phase relative to the box. Similarly, at the trailing edge, the applied delta V is decreased, thus decreasing the altitude and period and sliding the satellite forward in phase.” (SMAD) Also this function of maintaining the satellite within each box can be done autonomously either on board the satellite or on the ground “if the orbit control system fails, the ground or the on board system will determine that the satellite is slowly drifting from its assigned slot and a warning can be issued with adequate time to fix the problem or implement a back-up before adverse consequences occur.” (SMAD) It is important to note that it is more important to maintain attitude control than it is to keep the satellite within a certain position in the constellation. “A satellite that loses attitude control will usually tumble and then lose the payload function, power on the solar arrays, and contact with the ground. It also may point sensitive instruments at the sun or have substantial thermal problems. Even a brief attitude control failure can destroy the mission.” (SMAD) The thrusters and reaction wheels on board STD will be provided as follows. There will be two types of thrusters low and high. The high thrusters will mainly only be used after launch vehicle separation while the low thrusters will be used for station keeping and three axis attitude stability. There will be four reaction wheels aboard STD that will provide torque to control and maneuver the attitude with a speed range of 6000 rpm. All four reaction wheels will be used to maintain three-axis stabilization but if one should fail then the remaining three will still be able to provide stability.
The Attitude Determination and Control Subsystem (ADCS) of a space vehicle stabilizes and orients it n the desired directions during the mission despite the external disturbance torques acting on it (Wertz, & Larson, 1999). Spacecraft attitude refers to the angular orientation of a defined body’s fixed coordinate frame with respect to a separately defined external frame (Griffin 2004). Attitude determination is the process of measuring spacecraft orientation (Griffin 2004). Attitude control implies a process usually occurring more or less continuously of returning the spacecraft to a desired orientation, given that the measurement reveals a discrepancy. Measurement and actuation errors will result from inexact execution of reorientation maneuvers, based on inexact measurements, as well as from disturbances internally or externally. The space craft can’t respond instantly to this and time is consumed during measuring an error and computing and applying correction. The ADCS implements error detection and correction processes to ensure optimal station keeping and attitude determination. An ADCS system should be designed to resist these torques either passively, by exploiting inherent inertia or magnetic properties to make the “disturbances” stabilizing and their effects tolerable, or actively, by sensing the resulting motion and applying corrective torques. In addition to rejecting disturbances the ADCS must re-orient the vehicle to re-point the payload, solar arrays, or antennas. External references must be used to determine the vehicles absolute attitude to include the Sun, Earth’s IR horizon, the local magnetic field direction and the stars. The ADCS must control the vehicle attitude during firing of large liquid or solid rocket motors. Once on station the payload pointing requirements usually dominate. Maneuvers may be necessary to track stationary and moving targets. Passive stabilization techniques take advantage of basic physical principles and naturally occurring forces by designing the spacecraft to enhance the effect of one force while reducing others. In effect, we use the previously analyzed disturbance torques to control the spacecraft, choosing a design to emphasize one and mitigate the others. To effect the three- axis determination requires two vectors that can be measured in the spacecraft body frame and have known values in the inertial reference frame. The key lies in the type of sensor used to effect the measurement rather than in the nature of the reference point. The ADCS interacts with at least six other subsystems: EPS, Thermal, Propulsion, Communication, Structures, and the Payload. The majority of the components of the ADCS will be radiation hardened and will have a cross strapping for redundancy management. The ADCS of STD shall meet specific requirements. Those requirements are listed below with an explanation of how they will be met. Requirement One – The ADCS shall use multiple earth sensors to maintain an earth centric attitude. STDs ADCS will use multiple sensors: Attitude determination combines all available sensor information in a Kalman filter. Sensors are used to sense the attitude change from the desired position. The ADCS will have a star tracking sensor, both a fine and course sun sensor and an earth Sensor. The Star Tracker Assembly (STA) will provide an inertial attitude reference for use in orbit transfer state and super-synchronous Operations. The STA is not radiation hardened due to the nature of its purpose, to track using the radiation from stars. STDs STA can track up to 5 stars at a time. The STA has a redundancy capability in that only one star tracker is required by Attitude Determination for operations. The star tracker can provide attitude information only when it has located identifiable stars. (Pisacane, p.266) This is why we decided to use the 7.8 degree field of view. This field of view is large enough for one star tracker to provide 3 axis stabilized attitude. A star tracker is more like a camera that can provide information on several stars at the same time. The term ‘tracker’ refers to the ability to provide continuous position updates as the star moves. (Brown, p.304) Star sensors suffer from interference from the sun, earth and other bright sources. (Wertz p.186) Since there is the potential for interference STD will have other sensors the collectively establish a good attitude determination. Fine Sun Sensor Assembly (FSA) will provides a two axis inertial reference when Spacecraft attitude places the Sun in the field of view of the Fine Sun Sensor Assembly. The FSA is going to be radiation hardened. There will be two Single axis sensors and electronics that are nadir pointing. These were chosen to measure yaw, while the earth sensor measures roll and pitch. The redundancy capability for the FSA is two fine sun sensor assemblies; if one fails the other will be used. Fine sun sensors use analog information and digital information. (Pisacane, p.261) Sun sensors determine the angles of the sun by determining which of the lights sensitive cells in the sensor is mostly illuminated.(Pisacane, p.261) Sun sensor measures solar array orientation with respect to the sun. (Pattan p.150) STD will also be equipped with a coarse sun sensor (CSS) that will provide a two axis inertial reference when Spacecraft attitude puts the Sun in the field of view. The CSS will be radiation hardened. The CSS is considered a passive device. The CSS is not an assembly has no electronics and needs no Spacecraft power. Like the solar arrays it uses photovoltaic devices when exposed too sun light, the face chips convert sun light into electric current. The redundancy is there are two CSSs once on station CSS2 is stowed until needed. The final sensor used within the ADCS of STD will be the Earth Sensor Assembly (ESA). The ESA will provide two-axis Earth nadir reference when spacecraft attitude places the earth in the field of view of the Earth Sensor Assembly. The ESA will be radiation hardened. The ESA will have a Static horizon Sensor to ensure the earth is constantly on the horizon. This sensor will not use scanning. It will use thermopile detectors sense the Earth. Earth sensors will be used during LEO operations to help the spacecraft establish a vector on the earth in reference to its final orbital determination. The Earth sensor will ensure the X axis is constantly locked on to the Earth. If the Earth sensor looses lock then we have an anomaly. Earth sensors will measure orientation of the spacecraft with respect to the earth. (Pattan, p. 150) Requirement Two – The ADCS shall utilize reaction wheels for minor movements or attitude adjustments. Reaction Wheel Assembly (RWA) provides torque to control attitude and maneuver to a new attitude. The RWA is going to store momentum due to solar pressure. A reaction wheel is a flywheel driven by a reversible dc motor. To perform a maneuver with the reaction wheel, the flywheel is accelerated. (Brown, p.176) The center of pressure is offset from the SV center of mass. Resulting solar torque, over time, results in stored momentum. The RWA will be radiation hardened. The RWA will have built in redundancy in the form of four reaction wheels to provide three axis stabilization. A minimum three 3 assemblies are needed for 3 axis control. In the event one wheel fails the three remaining provide three axis torques. Torque capability and momentum storage is reduced. After turn off of one, RWA control performance will be degraded until a database is upgraded consistent with the remaining three. The RWAs on STD will provide linear smooth torque for attitude control. RWA control avoids excitation of significant space vehicle structural modes. Use of a reaction wheel on a spacecraft offers the advantage of substantial gyroscopic stability. RWA are generally used on spacecraft that require a relatively consistent pointing direction. RWA offers the advantages of high-precision, independent control about all three axes. Requirement Three – Magnetic torquers shall be used for momentum dumping of reaction wheels. Reaction wheels are used in spacecraft attitude control systems to stabilize the attitude of the spacecraft. Various external torques increase the spacecrafts angular momentum and speed up the reaction wheels. The stored momentum needs to be removed. STD will have a built in over speed protection – if RWA speed reaches the over speed limit of (+or -7000 rpm) – Magnetic torquers shall be used for momentum dumping of reaction wheels
The propellant budget for the STD is broken down into several categories. The categories include velocity correction and control, attitude control, nominal propellant, margin, residual, and finally, total propellant. The propellant required for velocity correction and control is estimated at 160Kg. The propellant required for attitude control is estimated at 125Kg. This equates to a nominal propellant requirement of 285Kg. A propellant margin of 10% is used with a residual of 1%. The overall propellant budget for STD is estimated at 320Kg. The propulsion system for the STD is broken down into several operational phases and components. The phases include launch, orbit transfer, station keeping, re-positioning, and attitude control. Propulsion requirements for launch are met with the appropriate choice of launch vehicle; the propulsion concept of operation follows: The selected launch vehicle will carry the required amount and type of propellant to achieve its initial orbit. Upon reaching the transfer orbit and separation, thruster manifold venting will occur, and both Rocket Engine Assembly (REA) and Liquid Apogee Engine (LAE) valves will open for fuel priming. Shortly after, fuel tank pressurization will occur followed by oxidizer tank pressurization. Pressurization will continue in preparation for apogee velocity burns. The final stage of the launch vehicle will remain attached to the spacecraft and will provide the propellant required by the LAE to reach the STD final operational orbit. Upon reaching operational orbit, station-keeping and re-positioning will be conducted using the STD Integrated Propulsion System (IPS), which includes four on-board thrusters and a Hydrazine fuel supply. The propulsion subsystem was selected based on the performance requirements as they relate to the operation and weight of the spacecraft. The launch vehicle propulsion system will not only launch the vehicle into an initial orbit, but must create enough controlled velocity to put it into its final super-synchronous orbit at an altitude of approximately 42,000 Km. The launch vehicle propellant required to reach the STD initial orbit is 10,000Kg. The propellant required in the final stage of the launch vehicle for reaching the STD final orbit is 3,000Kg. Once in its final orbit, station-keeping and re-positioning operations will require 150Kg of propellant and be handled by the spacecraft propulsion subsystem. There is not a propulsion requirement for de-saturation or attitude control, as those operations will be conducted primarily by on-board magnetic torquers and reaction wheels respectively. This being stated, propulsion can be used as a back-up if problems or failures occur within the primary system. As the spacecraft is already operating in super-synchronous orbit, there is no propellant dedicated for end of life maneuvers. The STD uses a hypergolic, bipropellant propulsion system. This system was chosen for its performance relative to a monopropellant system. The system will utilize a combination of Mono Methyl Hydrazine (MMH), Nitrogen Tetroxide (NTO), and Helium (He). Hydrazine was chosen as the fuel source based on analysis with other commonly used and readily available liquid fuel sources. MMH was chosen based on its performance and stability. NTO was chosen over nitric acid, another common oxidizer, as it is less corrosive. Propellant on-board the STD is pressurized using Helium gas. This system was chosen over pump or motor based systems for its simplicity and weight savings on-board the spacecraft. This simplifies the storage system by alleviating the management and added components required to manage liquid propellants. Helium gas was selected based on its availability and proven performance history. In addition to fuel types used for operations, the STD propulsion system uses several types of hardware components. The components include propellant tanks and manifolds, thrusters, valves, filters, and wiring harnesses. The STD utilizes a fuel tank, two oxidizer tanks, and two pressure tanks. For fuel storage, several tank designs were analyzed. Bulkhead and nested tanks were considered to reduce horizontal space requirements but the vertical distance required to contain the required fuel quantity exceeded allowable space. This conclusion resulted in the selection of separate fuel and oxidizer tanks. The fuel, oxidizer, and pressurant tanks are manufactured from Titanium. This material is compatible with the chemicals used (Brown, 1996, p.115). The fuel tank measures 48″x 60″ and holds approximately 1900L or 1500Kg of Hydrazine. The twin oxidizer tanks each measure 24″x 44″ and hold approximately 345L or 500Kg of NTO per tank. Both the fuel and oxidizer tanks will utilize internal Titanium sponge and vane assemblies to maintain fuel integrity during maneuvers in coordination with pressurized gas. The twin pressurant tanks each measure 12″x 36″ and hold approximately 5000cc of Helium per tank. Propulsion manifolds for the fuel and oxidizer tanks are designed to meet the Maximum Expected Operating Pressure (MEOP). The MEOP for the STD fuel and oxidizer system is rated at 300 PSIG. Burst PSIG uses a safety factor of four times for tubes and fittings and 2.5 times on manifold components. The MEOP for the STD pressurization tank is 4500 under high pressure and 300 under low pressure. Burst PSIG uses a safety factor of four times for tubes, fittings, and components. The STD uses one LAE and four on-board REAs for station-keeping and re-positioning. The LAE will be used to perform orbit raising maneuvers. The LAE is located on the -Z axis is rated at 120 lbs and draws fuel from the final stage Hydrazine tank. The nozzle will be made from Columbium and the LAE will produce a specific impulse greater than 325 seconds. The STD will also use an Oxidizer Depletion Detection (ODD) system. This system will shut down the LAE once the oxidizer tank is empty. The purpose of this system is to control Hydrazine consumption and to protect the spacecraft. This system monitors the LAE burn electronically and acts autonomously when pre-calculated thresholds are exceeded. The four REAs are rated at 2Kw and 1lbf and are located on the -/+ Y and -/+Z axes. The REAs each require a power cable and a power conditioning unit that will tie into the central propulsion wiring harness. Valves are located throughout the manifold systems to ensure proper fuel / oxidizer separation and to ensure propellant management is optimized. Several types of valves are used within the STD propellant management system, to include check, pyro, and latch valves. Check valves are used in several locations to prevent fuel and oxidizer vapor mixing. Pyro valves are used to stop or start fluid movement and were chosen for reliability and simplicity. Latch valves are used to isolate fuel to the thrusters. Propellant filters are used in line between the propellant tanks and thrusters to provide added protection against contamination. Thruster commands are routed through a central wiring harness which utilizes a single fault tolerant electrical design. This design incorporates redundant valve drivers and coils on the LAE and all REAs. Propulsion system components either operate mechanically or upon command. The LAE and REA valves require a command to turn either ON or OFF to initiate / shut off thrust. The pyro valves require a command to either open or close. The latch valves require a command to either open or close, depending on what configuration the valve is in prior to the command. All other valves on the STD propulsion system operate mechanically. The propulsion system will interface with the thermal, structure and mechanisms, electrical power, TT&C, and the ADC subsystems of the STD. The thermal subsystem will integrate tank heaters, line heaters, and valve heaters to ensure the propellant and its operating systems are regulated at the required temperatures. Major thermal requirements include preventing the propellant from freezing by controlling valve, tank, and line heaters, and limiting thruster cold starts by maintaining required cat bed temperatures. The structure and mechanisms subsystem will integrate the proper structures and mounts for the fuel, oxidizer, and pressurization tanks. Titanium is the primary material used for the propulsion system components and all structural materials associated with it must be compatible. This includes brackets for the manifolds, lines, valves, and wiring harnesses. The electrical power subsystem will ensure that all valve, heater, thermistor, and transducer power requirements pertaining to propulsion are met. This also includes ensuring the wiring harness meets all voltage and load requirements. The TT&C subsystem will be integrated to ensure that all communications are in place for state of health monitoring and commanding the thrusters. The ADC subsystem will integrate to ensure that enough propellant is budgeted to provide attitude control during maneuvering and thrusting. The ADC system will minimize propellant usage and maximize thruster life.
STD shall have 4 dedicated ground sites (See Appendix 2). All four dedicated ground sites will be equipped for carrier tracking, command reception and detection, telemetry modulation and transmission, ranging, and subsystem operations. These functions are basic to most space systems. However, with respect to STD, each of these requirements will be explained in this section. STD carrier tracking shall cover three specific functions: two-way coherent communication, two-way non-coherent communication, and 1-way communication. Two-way coherent communication is when the downlink carrier is transmitted so that its phase “synchronizes with the received phase of the uplink carrier” (Wertz, & Larson, 1999). This will allow for the downlink carrier frequency to be precisely offset from the uplink frequency and ensure the frequencies don’t interfere with each other. At all times, the downlink frequency will have a constant phase difference from the uplink. Two-way non-coherent communication is defined as, “transmission of information which does not make use of the carrier derived in the up-link receiver” (Lindsey, 1996). In other words, two-way non-coherent communication is simply a manual separation of the frequency and/or phase of the uplink and downlink. STD will employ this method as a back-up to coherent two-way communication. In the event the uplink or downlink hardware is damaged, non-coherent communication shall be utilized. Furthermore, this implies that STD will employ a one-way non-coherent means of communication as well. The one-way means of communication refers to the ability to operate a manually programmed single frequency or phase modulation on an uplink or downlink (Lindsey, 1996). STD command reception and detection shall include the ability to acquire/ track an uplink carrier, demodulation of carrier and subcarrier, derivation of bit timing and detection of data bits, resolving data-phase ambiguity when it exists, and forwarding of command data, clock, and in-lock indicator to the subsystem for command and data handling (Wertz, & Larson, 1999). Acquisition and tracking capabilities of STD satellite TT&C components shall be equipped with comparable components as the dedicated ground antennas. STD antennas will automatically acquire and track the uplink signal as soon as the satellite is radiated by an uplink signal. Once STD has acquired and is actively tracking an uplink signal, it shall have the capability to demodulate the carrier and subcarrier waves and send the information to an on-board computer processing unit. The TT&C components shall then be equipped to derive any bit timing and detect imperfections in the signal. All inaccuracies in the signal derived from this process shall be corrected by TT&C components before sending the data to the main satellite bus. Lastly, TT&C components shall have the capability to forward all command data to the main satellite bus for processing. Telemetry modulation and transmission is another function that shall be executed by the TT&C components. This will entail a reception function from the command and data-handling subsystem, modulation of the downlink, and a composite transmission. From the command and data-handling subsystem, data will be sent to the TT&C components for preparation for downlink. Next, the data will be modulated to carrier and/or subcarrier waves onboard the satellite. After modulation and multiplexing (if necessary), TT&C components shall transmit the downlink signal to a ground antenna (Wertz, & Larson, 1999). Ranging and subsystem operations are other specific functions to be carried out by TT&C components. With respect to ranging, TT&C components shall detect a pseudorandom code found in the uplink and immediately transmit that same code in the downlink signal. This will be utilized for ephemeris and satellite tracking. With respect to subsystem operations, the TT&C subsystem must be integrated with the entire satellite bus for command and control purposes. Data containing satellite commands must be relayed to this subsystem, while the subsystem shall be equipped to respond to such commands. Furthermore, TT&C shall provide health status and subsystem telemetry to the command and data handling subsystem for downlink of satellite status as well as the performance of any preprogrammed mission command sequencing (Wertz, & Larson, 1999). The last specific functions of the TT&C subsystem deal with anomalous conditions in which an STD satellite might be found. For cases involving a tumbling spacecraft, loss of earth scenario, or anomalous attitude is found, the TT&C components shall be equipped with the capability to detect such scenario and transfer mission uplink & downlink operations to an omni-directional antenna. Via an omni-directional antenna, TT&C components shall receive uplink, demodulate data, transfer data to the bus, receive data from the bus, modulate data, and transmit information on a downlink signal just as nominal TT&C components (Wertz, & Larson, 1999).
The command and data handling (CD&H) unit on board the spacecraft shall be the sole point for spacecraft commanding. This unit will receive commanding data from the TT&C unit and allocate the commands to the proper subsystem. The CD&H subsystem will be fully integrated with every other subsystem for this reason. Commands will be sent to the appropriate subsystem for processing. Once the commands have been processed by the appropriate subsystems, the CD&H subsystem will await for command verification from the subsystem. Once command verification is obtained, it shall be sent to the TT&C subsystem for downlink. For nominal operations, the CD&H subsystem will route all telemetry to the TT&C subsystem for downlink during state of health checks (Wertz, & Larson, 1999).
The Electrical Power Subsystem (EPS) is the most important subsystem on the spacecraft. The EPS must power the payload and all the other subsystems of the spacecraft. The EPS can impose major limitations if not designed correctly. Battery storage capability and capacity is extremely important for times of eclipse. Like the majority of unmanned spacecraft, STD will use a combination of solar panels and batteries to ensure maximum spacecraft design life. The EPS on STD will provide optimal power for the life spacecraft and its mission. There are two peak times when the Electrical Power Subsystem needs to provide more power than day to day operations, they are the beginning of life and the end of life. Spacecraft design teams focus on four areas to account for these peak times and ensure that there is enough power stored within the batteries before launch and that there is plenty of power when the spacecraft is nearing the end of its life. The Electrical Power Subsystem (EPS) on STD will focus on those four main areas; power source, energy storage, power distribution and power regulation and control. The EPS for STD must meet four specific requirements. The following sections will discuss how those requirements are meet. Requirement One – STD shall be equipped with the most efficient solar array cells. Unmanned spacecraft requires reliable continuous operation of the power system to sustain vehicle through the mission. If the power system fails the spacecraft mission fails. Most of the other system components have a redundant capability or can be fixed from the ground. STD will rely on a battery backup if the solar arrays were to fail. But that backup capability will only last so long because the solar arrays are needed to recharge the batteries. To provide the reliability and assurance that the spacecraft will have sufficient power throughout its life span, STD will have two 5000 watt solar arrays with Photovoltaic solar cells as the primary source of power. Each solar array wing will have four panels a solar array boom and harness and deployment hinges. Each of the solar arrays will be driven by a Solar Array Drive Assembly (SADA) attached to the positive and negative structures of the spacecraft. The SADA will be connected to the Sun sensor. The direct link to the sun sensor will allow the solar array to be able to track and maintain locked on the sun for the longest periods of time. The Sun sensor also supports the ADCS in keeping the spacecraft in the correct orientation. A solar array’s illumination intensity depends on orbital parameters such as the sun angles, eclipse periods, solar distance, and concentration of solar energy (Wertz, & Larson, 1999). One of the main design hurdles in designing solar arrays is to realistically determine the needed power production to support our mission. Each of the solar arrays will have four panels of Photovoltaic cells. Photovoltaic cells are optimal for super-synchronous orbits because they are reliable. This type of solar cell will convert solar radiation directly into electrical power (Wertz, & Larson, 1999). Each solar cell on the solar array will convert the oxidation reaction into electricity. These cells are more tolerant to the radiation in space. They also give greater output power efficiency, and are more heat tolerant. With every advantage come disadvantages. These solar cells are heavier and more fragile. So that makes them more costly to produce. (Pattan p.175) Since the power from individual cells is actually very low. In order to meet the power needs the cells are placed on panels in a series of combinations. These combinations will help the solar arrays provide the most efficient to the spacecraft and it subsystems. Requirement Two – STD will have secondary (rechargeable) batteries Energy storage is just as important as gathering the energy. When a spacecraft goes into periods of eclipse the batteries will be required to maintain the spacecraft during the period when the solar arrays do not see the sun. A spacecraft battery consists of individual cells connected in series. The actual number of cells on the spacecraft will depend on the overall needed bus voltage. Batteries can be connected to increase in watt-hour capacity (Wertz, & Larson, 1999). STD will use the secondary battery type instead of the primary batteries. A secondary battery has electrodes that can be reconstituted by passing electricity back through it; also called a storage or rechargeable battery, it can be reused many times. STD will use batteries as a place to store electricity in a chemical form until it is needed to power the space craft during eclipse. The batteries will provide a constant source of power to the spacecraft. To store the energy created from the power the solar arrays collect STD will use completely redundant battery system with 36 separate pressure vessels. Each one is acting as a single battery, but they are cross strapped to ensure redundant capability. The batteries are external to the spacecraft structures. Each vessel sends temp, pressure and voltage readings from each individual vessel. STD will use Lithium Ion Batteries. The rechargeable Lithium Ion battery is the latest technology for space craft. They are easy to package and light; this will reduce the weight of the EPS and leave room for other components or fuel. Requirement three – The EPS on STD shall distribute power throughout the spacecraft depending on load requirements, and subsystem functions. Spacecraft power distribution consists of cabling, fault protection, and switches to turn power on and off. STD will use mechanical power switches. Power converters will be attached as the interface between each of the spacecraft components. These converters will ensure that the correct amount of power reaches the components and does not over load the circuits. Properly ensuring that the high voltage lines go to the high voltage components and the low voltage lines go to the low voltage components. To reduce Electromagnetic-interference an ac-dc converter will be connected to the bus. The use of harnesses will help the wires from the batteries and solar arrays travel the shortest distance possible through the spacecraft structure. to reduce voltage drops. Like most spacecraft STD will use dc power distribution to assist in generating direct current. To keep the weight of the EPS as low as it can be the design team will avoid the use of ac conversion units. STD will incorporate a fault protection system to isolate errors that could potentially cause mission degradation or spacecraft loss. Requirement four – The EPS on STD shall regulate and control the power utilized by the spacecraft. To regulate the power on STD the amount of power the solar arrays generate, the EPS needs to be controlled to prevent battery over charging and excessive heating. STD will use louvers to dissipate extra power into the structure of the spacecraft to help assist in heating components. STD will use quasi-regulated bus voltage to regulate the amount of power stored in the batteries during the charging periods. This method fixes at a potential above the normal bus voltage during battery charging. The solar arrays will charge the batteries when the vehicle is not in the eclipse season. The batteries will be charged individually in a predefined sequence. Individually charging the batteries will reduce degradation and ensure the longest battery life.
Thermal components aboard the STD must take into account the radiation from the Sun and energy transmitted from the Earth. Components must be analyzed to account for energy absorption and release along with the resident structures within and without the spacecraft to determine the total energy balance (Gilmore, 2002, p. 21). This balance is designed to ensure all operational components maintain the required temperatures in order to operate effectively throughout the life of the mission. The primary subsystems and associated components that must be thermally regulated include the payload, propulsion, structures & mechanisms, ADC, TT&C, and EPS. TCS management will differ for each subsystem based on its required temperature range and operating environment. Table 1 defines the operational and survivable temperature ranges for the STD Space segment components. The TCS on-board STD uses passive and active controls to maintain the required operational and survival temperature ranges. Passive controls consist of a combination of Multilayer Insulation (MLI) blankets, surface finishes, and radiators. Active controls consist of heaters, louvers, and heat pipes. Thermal control software will be utilized to command and control active thermal components. Thermal controls are assigned to each subsystem. The STD will operate in several thermal environments. During ascent, when the payload is stowed, it will be exposed to combination of molecular heating and rotisserie. When the spacecraft is in its transfer orbit, it will be exposed to rotisserie and Earth eclipse, along with LAE firings. When undergoing standard mission operations in its operational or final orbit, the STD will be exposed to a combination of radiation, eclipses, and solstices. These conditions will cause great fluctuations in operating temperatures on-board the spacecraft. The overall external surface of the STD will be covered in MLI. Cut-outs in the MLI are integrated into the blanket to provide a means to release heat when required. All subsystems are thermally protected on the STD to ensure environmental damage does not occur. The battery on-board the STD is an individual pressure vessel Nickel-Hydrogen battery. To protect the battery, active thermal controls are used and consist of several software controlled heaters. The heaters operate from closed-loop temperature feedback and use primary and redundant heater strips. The heaters will prevent the battery from exceeding lower range temperatures. Thermistors are emplaced to provide temperature readings. Power for the heaters originates from the auxiliary power unit. Passive controls are also used and consist of a radiator panel and MLI. The thermal radiator faces to the outside of the spacecraft and is coated with a thermal control film. The radiator panel will prevent the battery from exceeding upper range temperatures. MLI covers all remaining unexposed surfaces of the battery. Excess power from the batteries will be shunted within the power bay to provide additional heat as required. The reaction wheels on the STD will reside in a thermally controlled bay. The bay will be insulated with MLI and the reaction wheel structure will mount to a heat plate with an imbedded Helium pipe to control reaction wheel temperatures. The bay will also contain two heaters and radiators for added thermal control. Earth, star, and sun sensors have different requirements based on function and spacecraft location. The Earth sensor will be isolated from the body of the spacecraft and will use MLI wrapped around the sensor body along with a heater and radiator to further control temperature. The STD uses a pair of star sensors. Each sensor will be cycled on and off every other hour to reduce heat and extend the life of the sensors. To protect the optical properties of the star sensors, the sensor will be isolated from its mounting bracket using thermal coatings, tape and MLI. A shutter will be used to protect the optics from direct radiation from the sun. A pair of sun sensors is used on the STD. The sun sensors will use a combination of thermal protective tape and MLI. Thermal control for the propulsion subsystem utilizes a combination of MLI, heaters, and thermal surface coatings. The propulsion tanks each have different temperature requirements. The oxidizer tanks each use a single heater, the fuel tank uses three heaters, and the pressurant tank uses one heater. All tanks use thermistors to manage temperature feedback. The LAE and REAs not only require thermal control for its propellant, but due to high temperatures during maneuvers, thermal control is required to protect the exterior of the spacecraft from thruster exhaust. To control temperatures, thruster bodies are isolated from the spacecraft using low conductivity materials and thermal coated heat shields are emplaced to protect the spacecraft from exhaust. The thruster bodies are wrapped in MLI and a combination of line heaters and radiators are used to manage thruster temperature. The propulsion manifolds and lines are coated with a high temperature coating and wrapped in MLI. Heaters are displaced along manifold lines where spacecraft density reduces temperatures thereby requiring active heating. The antenna requires several thermal applications. The feed horn and aperture will be covered with MLI. A thermal surface coating will be applied the front-side of the reflector and MLI will be used to protect the back-side. Heaters are attached to the antenna pointing mechanism and boom structure where hinges and mechanisms exist. Heaters will be command activated and used when required. Thermal coatings and tape are applied to subcomponents within those structures. There are four solar panels on the STD, two on each side of the spacecraft. To optimize current flow, the solar panel assembly must be kept cool. The solar cells themselves cannot be coated. All areas surrounding the solar cells are covered with either thermal coatings or tape. The solar booms are covered in thermal tape. Heaters are positioned to provide temperature control both prior to and after deployment. The telescope and focal plane assembly is the primary payload on the STD and requires the closest tolerances in regards to thermal protection. Any thermal distortion applied to the optics mirror and structure will have a highly negative impact on STD pointing accuracy and resolution. To protect the telescope and focal plane assembly, thermal coatings are used in combination with heat coils. The heat coils use Helium gas to add or subtract heat in coordination with radiators and solar panels. Heater size is reduced in order to mount a higher quantity of heaters in a given space. This allows more accurate, localized heat management throughout the structure. Mechanical attachments will be isolated using low-conductivity materials. The payload electronics compartment will be isolated towards the center of the payload to protect it from radiation. The electronics module with data processing and handling will mount to a heat plate with integral helium heat tubes. The heat tubes will be augmented with heaters and radiators to keep the thermal properties within operational limits. Thermistors will be used to provide feedback to the auxiliary heaters.
The main mission is to identify and track all objects within the geosynchronous belt. Accurate positional data will be relayed on to the Joint Space Operations Center for orbital conjunction assessment. Mapping of the entire geosynchronous belt will provide all users of data a mechanism for planning orbital maneuvers in addition to satellite positioning.
For the payload portion of the STD satellite we choose to utilize the technology developed and used for Space Based Infrared System (SBIRS) Highly Elliptical Orbit (HEO) payload. This technology will provide more than capable ability to detect and track objects throughout the Geosynchronous Equatorial Orbit (GEO) belt. Only a few minor modifications are needed to support the overall mission. The payload portion of the vehicle will rely on the bus portion for support i.e. communications, power, and data computation and storing. The payload portion will consist of six different subsystems contained within. They are the Telescope Assembly (TA), Focal Plane Assembly (FPA), Attitude Control Assembly (ACA), Processing Assembly (PA), Thermal Control Assembly (TCA), and Power & Signal Distribution Vehicle Assembly (PSA).
The main function of the TA is to house the FPA, collect, filter, focus energy onto the FPA and provide a controlled thermal environment for the collection of data. The open end of the telescope itself there will be an aperture door to protect the lenses from contamination, condensation and small micro particles that could affect the operation of the telescope. On the inner portion of the sun shade there will be a series of baffles to collect stray Infrared (IR) energy so the sensor will not be drowned out by the extra energy. By eliminating most of the stray IR signals the energy within the direct field of regard of the vehicle will be the only one processed. As the IR energy is able to enter directly through the sunshade it is then collected by one of two lenses from within the telescope. The first lens is known as a fisheye lens and has a very wide field of view (approximately 20kilometer field of view of the GEO belt. A second lens takes the energy collected, acts like a corrector lens and removes any remaining aberrations of the waveform so the energy can be focused directly onto a mirror series of two lenses. The mirror assembly will consist of a convex mirror design and have very high reflective properties. This mirror will focus the energy collected by the two lenses and focus it onto one of two focal plane assemblies. The OTA will also consist of two internal calibration bulbs found behind the main mirror assembly unit. The calibration assembly will be activated through ground commanding, the energy is transmitted through a small hole in the center of the main mirror assembly. The lamp will fully illuminate the entire focal plane array for calibration. These lamps will generate a visible light signal and a Long Wave IR (LWIR) signal. The function of this section is to ensure elements within the focal plane array are functioning as they are intended to. If there is a catastrophic failure of either calibration bulb a known source from a ground location can be used to measure intensity level. The internal calibration bulbs will have an overall lifespan of around 2000 hours, which should provide more than enough calibration for the lifetime of the vehicle. The more accurate the sensor the more accurately we can determine the exact signature of the object. Additional baffles are utilized between the main mirror assembly and the FPA. If calibration signals are collected from external sources such as a known intensity from a star, this information is interfaced down into the Attitude Control Assembly so calculations can be performed to maintain accurate pointing angles. This interface also interacts with the azimuth and elevation drive motors to keep the sensor pointing accurate. Onboard the satellite there is also a payload coordinate system specific to the vehicle to maintain accurate pointing of the sensor.
The purpose of the FPA is to have a certain spectral response to different intensity of light and IR signals and convert this energy into electrical signals to be sent for further processing. The success of the FPA functioning correctly is dependent on how cool the FPA is kept. This will be discussed further in the TCA. The FPA will consist of six separate Sensor Chip Assemblies (SCA’s) visible arrays and two LWIR arrays. Each array will consist of the latest in technology at the time of this writing 2048X2048 4 quadrant staring arrays; the arrays will also filter out all remaining waveforms other than LWIR and Visible. The FPA will produce a resolution of better than 1 cm resolution at an altitude of 314 km. The visible arrays will give the advantage of target identification with very low magnitude light waves being detected. Light reflections will be picked up due to the sun behind the sensor. This sensor is designed to operate only during daylight hours even though it has the capability of nighttime operation by utilizing the LWIR sensors. The FPA will be suspended out to the middle of the TA. One disadvantage to this is the arms holding the FPA out to the center of the TA will block a certain amount of energy allowed to enter the telescope. The area blocked will be filtered out with the processing of the data. Also contained within the FPA is an Analog Pre Processor (APP), this device takes the analog signal given off the FPA, amplifies, and performs the initial processing/filtering of the data. The data is also digitized before it leaves the FPA. This is where we take certain exceedance levels and establish them so that only the target data is processed and all other is converted into heat and utilized to help maintain vehicle temperature. An interface with the TCA will also determine FPA temperature so heating or cooling can take place to maintain optimal FPA temperature. A series of cables will act as an interface between the FPA and other subsystems on the vehicle.
The function of the ACA is to provide control and knowledge of the actual sensors position in relation to the vehicle itself as well as known target sets (i.e. stars). The processing of positional data is separated into onboard and ground functions. Ground processors identify the targets and the onboard processors move the sensor to the target if requested. There are a number of different effects on the sensor by which the sensor might have to be moved over time to accurately support the mission. The mission portion will house a gyro assembly consisting of 4 separate gyros to indicate movement of the sensor in multiple directions. The gyro assembly unit will be redundant with the addition of a fourth gyro in the event one fails to operate. Positional data is collected and sent into the ACA for processing so commands may be sent to control the position of the sensor. Commands come from the ground station and are executed to move the elevation drive motor for a change in elevation angle, while the commands executed to move the azimuth drive motor are used to change azimuth. On board flight software will interface with the ACA, especially during the prediction of sun directly affecting the senor. A command will be internally generated by the onboard flight software to close the sun shade. But most of this processing is a result of the satellites position collected by the ACA.
The function of the PA is to process all of the mission data, control all internal calibration commanding and processing of all state of health data collected from various transducers throughout the vehicle. There will be two onboard computers providing total redundancy in the event of a failure. The processing of mission data as well as all commanding data and timing are all handled by this assembly. Attitude control signals are also processed within for pointing. The majority of the processing comes from the IR and Visible data compression. Matching data up with onboard solution sets (algorithms) to help reduce noise and clutter helps with the initial target identification. The majority of data processing will take place from the ground facility. The PA formats both telemetry and mission data into packets for transmission down to the ground station. On board buffers will store information until recorded by the onboard recording device.
The purpose of the payload TCA is to maintain all thermal sensitive components of the sensor payload, within its desired temperature range, for all environmental and operational conditions. TCA can be divided into three distinct sections; electronics, mechanisms and all others. For the electronics section this is where the cryogenic radiator dissipates heat coming from the focusing of energy on the FPA as well s the telescope. The operating temperature of the LWIR SCA’s require it to maintain a temperature of <126k and to do this it must be cryogenically cooled. The radiator itself will contain a second surface mirror material to provide cooling. The mechanisms section, the TCA provides and maintains temperature through various means, thermal coatings, reflective films, reflective insulation, kapton film, isolation material. Also small electrical heaters (thermistors) are used to keep mechanical/component temperature to a minimum during a survival mode of operation. The last portion includes the gyro assembly unit, health and status unit, and analog pre processor. The orbital position of the satellite will often times determine whether or not a certain amount of heating or cooling needs to take place to maintain on orbit status in support of the mission.
The purpose of the PSA is to provide power to the payload to satisfy the mission requirements set forth. This assembly consists of all power supplies, cabling and related components needed to distribute power and signal to the other payload subsystems. This subsystem shall receive both primary and backup power from the host space vehicle. The power will be divided into three separate sections; first distribution of power goes to the line of sight computer and all the TSA associated hardware. The second section is the main processor to perform the calculations and procession in support of mission data. Third area, power is further distributed to the drive assembly motors to change position of the sensor. The last area power is distributed into is the TCA for survival heaters. There are two sides A and B for redundant power coming into the mission portion.
The ground segment of a space system is vital to system operations and mission success. The ground segment must adequately support the space segment (i.e., satellites) for a system. Ground stations and equipment are used to command and control (C2) the constellation, monitor the status of health of the vehicles via telemetry data transmitted from spacecraft subsystems, and track the positions of the spacecraft. Ground systems must also receive mission data from the satellites and distribute to the applicable users. In the case of STD, the ground segment must perform all of these functions. Without an adequate supporting ground system, the data collected by and transmitted from STD is useless. Obviously the ground segment is responsible for commanding and controlling the satellites. The positioning of the STD satellites must be monitored and tracked precisely to ensure appropriate operations. In order to successfully contribute to mission success, a ground segment must satisfy specific requirements. For example, the data shall be transmitted in the appropriate formats and able to be translated into a useable form for users. Depending on the desired end user equipment and desired results, this may require that data be formatted for human-readable, binary, hexadecimal, or many other computer languages. For STD, imagery must be displayed properly, and at the required level of clarity, and any accompanying information must be human-readable. Whenever Dept. of Defense (DoD) assets are involved, there are often also additional requirements to satisfy, such as security and classification standards. Due to the sensitive nature of many satellites STD will be viewing and monitoring in geosynchronous belt, security is certainly a consideration. The secure integrity of the data shall be maintained throughout the communications transmission links. The signal will be encoded on the downlink to protect user mission data information and will be encoded on the uplink as the system will be managed by the DoD. Because they are so complex in nature, satellite constellations rely heavily on intricate and elaborate networks. The C2 of the STD satellites must have well-defined, effective ground segment equipment architecture. The Colorado Mission Control (CMC) Station shall reside in Colorado Springs, Colorado. Much of the necessary expertise will reside directly on the MCC Operations Floor, led by the Mission Commander (MC). The Multi-Mission Planner (MMP) and Multi-Mission Operator (MMO) will plan and execute mission maneuvers, respectively, while the Orbital Analysis (OA) personnel plan, monitor, and ensure the satellites’ orbits in space as well as Collision Avoidance (COLA). The Satellite Systems Operator (SSO) and Ground Systems Operator (GSO) will be supported by a civilian Satellite System Engineer (SSE) and Ground System Engineer (GSE), all of whom work under the Systems Crew Chief (SC). In addition to the Operations Floor, the MCC will comprise a Conference Room (with Operations Floor viewing area), and Anomaly Resolution Room. Another requirement is that the satellites shall obviously have visibility to and from the required ground stations. This is important not only for sending C2 data and receiving telemetry data but also for distributing mission data to the end users. The amount of satellite visibility to ground stations can hinge on the amount of data storage onboard each satellite as well as the frequency with which user data must be sent. STD will incorporate three dedicated sites in addition to the primary MCC: one in Alice Springs, Australia; one on Diego Garcia; and one at Fylingdales AB, United Kingdom). For additional redundancy purposes, STD shall be able to utilize the eight sites of the Air Force Satellite Control Network (AFSCN). In order to do this, certain stipulations must be taken into consideration. The antennas, signals, data rates, frequency, and other characteristics of the transmission links must all be compatible with those of the AFSCN. Specific data rates and frequency ranges must be applied to guarantee data distribution and security requirements are met. The STD will utilize the S-band (2025-2120 MHz) and X-band (7145-7235 MHz) spectrums for uplink communications frequencies, and the S-band (2200-2300 MHz) for downlink frequencies. The minimum acceptable level of Effective Isotropic Radiated Power (EIRP) shall be 74 dB and the acceptable level for Gain/Temperature (G/T) shall be 29 dB for S-band. The standard data rate for the transmission links will be no more than 105 Mbs, with the telemetry stream being limited to approximately 1 Mbs. These data rates, along with other signal characteristics, must guarantee compatibility not only between the ground terminals and satellites, but also with the AFSCN Space-Ground Link System (SGLS), for anomalous C2 purposes. All of the involved antennas, in order to communicate, must be able to transmit and receive the same waveforms in the same frequencies using the same modulation and encryption schemes. For an added layer of data security, the STD will employ a passband direct sequence spread spectrum (DSSS) modulation method, adding a continuous stream of pseudo-randomly-generated noise to the signal. Because of this, uplink and downlink sequences for all antennas must be synchronized. The ground segment, being such an integral part of the overall system, interfaces with many of the other segments and subsystems of the program. A ground segment obviously interfaces extensively with the space segment, particularly the communications subsystem. This ground-to-space/space-to-ground link is the key to all communications, both C2 and mission-related, to and from the satellites; thus, the success of the mission and the health of the constellation depend on it functioning properly. In addition to the subsystems of the space segment, the ground segment has many internal interfaces with its various subsystems. Relay ground station interaction may be required for the data to be distributed to other parts of the world (although this is not the case with STD due to its multiple ground sites and alternate AFSCN capability). Another link vital to successful mission execution is that which connects the MCC and the end users. These networks can take on many different configurations. The STD program will utilize an architecture in which the MCC will receive the mission data, process the imagery, and then distribute the files to customers via terrestrial networks. The STD ground segment can be further divided into subsystems. The antennas themselves, used for commanding the satellites and receiving telemetry and data streams, will make use of 120″ dishes, with two antennas located at the CMC Station as well as one at each of the three supporting locations. The mission data encompasses imagery taken by the STD optical sensor, as well as other corresponding identification and descriptive support data particular to the satellite being photographed. The users of the data will be the participating organizational stakeholders who own and operate satellites in the geosynchronous belt as well as various U.S. space tracking agencies. A terrestrial communications network will link all ground stations to the CMC Station, so that both mission data and command/telemetry data may be distributed to the facility. Once received by the CMC, the mission data will be stored on a centralized secure server and backed-up on a daily basis to help mitigate losses due to potential outages. In order to make the mission data to accessible to users, they will be required to register for accounts. The accounts will grant each organization access to the mission data for its respective satellite constellation(s). U.S. space tracking agencies will be granted access to all mission data.
The STD spacecraft will require a heavy lift capable launch vehicle. Two launch vehicles were considered to meeting the requirements to put STD into a super-synchronous orbit. The Russian Proton launch vehicle launched from the Baikonur Cosmodrome and the Atlas V. The Proton launch vehicle provides a low cost alternative, while meeting strict launch date timelines. Even though maintaining low costs are a priority the reliability of this vehicle is not as good as the Atlas. After careful consideration the best solution would be to proceed with the Atlas V. Selection for the Atlas V was based on the overall weight of payload, spacecraft dimensions, final orbit and mission timelines. Based on all of the above the Atlas V was selected due to its heavy lift capability not offered by many other manufacturers. The core Atlas vehicle is known for its capability and reliability in placing spacecraft into a geosynchronous orbit. Since the first Atlas V launch in August 2002, nearly 20 launches, it has enjoyed a near perfect success rate, the only anomaly was not with the primary vehicle itself but instead with the Centaur upper stage. After nearly 52 years of venerable service to the United States and the world’s space endeavors, the veteran Atlas rocket vehicle has embarked on its 600th mission Sunday, launching the USAF Defense Meteorological Satellite Program F18 (DMSP F18) satellite from Vandenberg AFB on the newest member of the Atlas rocket family: the Atlas V.. (Chris Gebhardt , 2009) One thing to add is the reliability of a launch vehicle will drive the cost of insurance for the customer down or lack thereof will raise the cost. The Atlas V will be launched from the Eastern Range, Cape Kennedy Space Center 22.30N 80.33W, launched due east. The launch services provided for the Atlas V include the following; the facilities in support of spacecraft integration, processing, encapsulation, launch operations and verification of specific orbital parameters. From the moment the spacecraft arrives in Florida, it will be sent to a payload processing facility where it is integrated with the adapter portion of the launch vehicle. At this time the payload fairing will be added to protect the vehicle as it makes its way out to the vertical integration facility. A 5.4 meter diameter payload fairing was selected to allow for some margin of capability when placing a satellite such as STD to Super-Synch orbit. This payload fairing is the largest manufactured for the selected launch vehicle. A list of mission requirements for special orbital maneuvers, specific orbital requirements and any additional tracking needed to maintain accurate ephemeris data. Another area that falls under mission requirements is any unique communications and data routing through the ground stations. It will be necessary for a link from the operations center providing command and control to the vehicle be established.
The Atlas V has the capability of placing between 10,470 – 28,660 lbs into GEO orbit. The Atlas V vehicle in conjunction with the onboard apogee engine will use a three-burn Geo-transfer orbit injection. This is the most efficient method of placing a spacecraft into a super-synch orbit. The Atlas V launch system has all the ground processing hardware and facility to meet the needs of the STD. The primary interface between the launch vehicle is the payload adapter ring connecting the Centaur upper stage with the spacecraft. This adapter provides the interface with the ground as well as the mechanisms required for separation from the launch vehicle. A primary control link will be established between the launch control facilities and the Colorado Mission Control Station (CMCS). Telemetry data will be shipped from the Centaur upper stage by using the Digital Telepak and Telemetry and Data Relay Satellite System (TDRSS) transmitter, where data is forwarded on to the CMCS station. Air Force Satellite Control Network (AFSCN) will be utilized during orbital injection for near continuous support. Specifically Diego Garcia will be used for initial contact during launch and quickly followed up by Guam Tracking Station. The observations provided from these tracking stations will feed into the overall updating of ephemeris data for the refinement of orbital determination. The launch site operations support provides all the prelaunch preparation of the payload adapter and all other spacecraft support hardware. Transportation of the encapsulated spacecraft and mating of the spacecraft with the launch vehicle takes place at the launch pad. The entire support of the launch vehicle and spacecraft interface tests fall under the launch site operations section. Facilities at the launch pad provide the spacecraft with power, instrumentation, water, gasses, temperature control and communications.
Mission scheduling for launches are typically planned 12-18 months in advance. A launch window will be identified from opening to closing and should be down to the nearest minute in accuracy. Normally the launch window will be propagated out for at least two weeks after the final closing. Because this is the primary payload and only payload the launch window will be constrained to the payload on board. The STD will have an Interface Requirements Document and this document will include all the interface details for all the on board systems.
All the special requirements set forth for the payload fairing as far as size, shape, weight information. Any additional access required to the spacecraft during the pre-launch checkout will be noted and access panels will be identified. Fairing temperatures will be maintained according to specifications while on the ground. There might be slight variations in the way the fairing supports and protects the payload from an internal perspective.
There will be a document identifying any and all testing to be performed while in pre-launch and this will identify static loads, vibration, acoustics, shock, and the results of all systems within the satellite for functionality. The functionality of these tests should confirm the readiness of the spacecraft for a space environment. Additionally a spacecraft structural load analysis report should be performed prior to launch.
Prior to the arrival of the payload the safety issues will be identified specific to the spacecraft and vehicle. This will identify any and all hazardous items such as exploding ordinance devices, propellants, pressurized system and any RF radiation given off.
Spacecraft mass must be identified at each stage through orbital insertion. Important factors are the center of balance and any shifts occurring during the deployment of any appendage. Fuel also has some considerations when it comes to mass with the slosh of that fuel during the different stages of insertion.
The crosscutting technical management portion of the STD system is what creates the product from cradle to grave along with all the planning involved. In order to turn STD into a fully functional product that meets the stakeholders needs “every member of the technical team relies on technical planning; management of requirements, interfaces, technical risk, configuration, and technical data; technical assessment; and decision analysis to meet the projects objectives.” (NASA) Once the planning steps have been established, “the technical team will generate a technical cost estimate and schedule based on the actual work required to satisfy the projects technical requirements.” (NASA) The cost and schedule estimate are important so that the program has a budget and an end goal. Even though, for STD our cost and schedule are unlimited, we should still use realistic values for cost and the schedule because it is important for a system engineer to report accurate cost and schedule estimates or the project could run over budget and behind schedule. “The current NASA policy is that projects are to submit budgets sufficient to ensure a 70 percent probability of achieving the objectives within the proposed resources.” (NASA) In order to provide an accurate schedule “a complete network schedule may be used to calculate how long it will take to complete a project; which activities determine the duration; and how much spare time exists for all other activities of the project.” (NASA) It is important to note that before the project goes any further “the technical team should ensure that the appropriate stakeholders have a method to provide inputs and review the project planning for implementation of stakeholders interest” (NASA) since it is for the stakeholders that STD is being built in the first place. During the system engineering process it is important to watch out for problems that might interrupt the project. One of these problems is requirement creep “requirement creep is the term used to describe the subtle way that the requirements grow imperceptibly during the course of a project. The tendency for the set of requirements is to relentlessly increase in size during the course of development, resulting in a system that is more expensive and complex than originally intended. Often the changes are quite innocent and what appear to be changes to a system are really enhancements in disguise.” (NASA) This increase in requirements is caused by having many stakeholders on one project as is the case with STD. Along with problems there are a few risk that must be taken into consideration like: cost, schedule, technical, and programmatic risk. Both cost and schedule risk are related because if the cost of the project increases this might increase the schedule while the project waits for additional funds, and if the schedule increased due to a delay then this would increase the cost to keep the project running longer. One way to reduce technical risk is through configuration management “configuration management reduces technical risk b ensuring the correct product configurations, distinguishes among product versions, ensures consistency between the product and information about the product, and avoids the embarrassment of stakeholder dissatisfaction and complaint.” (NASA) Programmatic risk is simply outside factors that affect the project that are out of the system engineers’ control. One possible outcome of programmatic risk is project termination “it should be noted that project termination, while usually disappointing to project personnel, may be a proper reaction to changes in external conditions” (NASA) Additional risk to the project could be inadequate staffing or skills, operational hazards, poorly defined requirements, infeasible design, unavailable technology, or low quality assurance. One way to reduce risk to the project is to have an alternative or back up plan, in the case of STD there is redundancy throughout the project from having four satellites to its extra reaction wheel. Having alternatives and backups designed into the system assures that the stakeholders are getting their monies worth. Before the STD system can be approved to start the engineering process, several reviews must be done. First, a mission concept review will be done to determine if the mission of STD fully satisfies the customer’s needs. Next will be the production readiness review which will “determine the readiness of the system developers to efficiently produce the required number of systems. It ensures that the production plans; fabrication, assembly, and integration enabling products; and personnel are in place and ready to begin production.” (NASA) After that is the flight readiness review which is mainly done by a computer program given all of the flight characteristics of STD and its probability to survive the launch. Next is the post launch assessment that will determine if all systems aboard STD are working properly and can carry out its full operations. Toward the end life of STD a decommissioning review of STD will be done to determine the best way to get rid of STD without it effecting other satellite orbits or possible collisions. Once the project has been completed it is important for the system engineer to compile lessons learned so that others will not make the same mistake. “A lesson learned is knowledge or understanding gained by experience-either a successful test or mission or a mishap or failure.” (NASA)
The Geo Belt is a unique place because it is the only orbit that can have a satellite maintain a constant position over the Earth. Because of this special ability over the years the Geo Belt has become a very crowded place therefore it is important to maintain situational awareness in this special region of space. As the years go by room in the Geo Belt will only increase in demand for new satellites. This demand emphasizes the importance of STD to accomplish its mission of situational awareness by tracking and mapping objects in the Geo Belt.
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