Do It Yourself Solar Energy

DIY Home Energy System

This ebook guide teaches you how to escape complete dependence the power grid and learn how to live mostly on your own power and make sure that you are dependent on Yourself. You will be able to slash your energy bill by over 75% and not have to depend on greedy energy companies. The largest energy corporations are a monopoly for a given area, so they do not need to care about customer service or doing right by the people they service. You will learn how to break this monopoly and depend on yourself. Make your home immune to power shortages, blackouts, and energy failures; live free of any worry that the grid will totally fail you! You will learn practical steps such as how to build your own solar panel for less than $60! Once you start relying more on solar power you will be able to easily protect your family from dangerous power outages, and live free! Continue reading...

DIY Home Energy System Summary

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Author: Jeff Davis
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My DIY Home Energy System Review

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Sail vs VTilt Solar Panels

On Mariner spacecraft the thousands of solar cells that convert sunlight into electrical energy were mounted on the face of flat rectangular panels extending like wings from the spacecraft. Since the new spacecraft had to travel from the orbit of the Earth to that of Mercury, its solar cell energy-gathering system had to accommodate to the change of nearly 5 times in the amount of solar radiation that would be received. Early studies by JPL and Boeing concluded that the best way to keep the solar panels at the right temperature of about 100 C (212 F), while still providing a fairly constant power output from them into the spacecraft electrical system and also meeting the weight constraints, would be to Fig. 4-1. One method of safeguarding the solar panels from overheating as Mariner approached closer to the Sun was to tilt them into a V-configuration. Fig. 4-1. One method of safeguarding the solar panels from overheating as Mariner approached closer to the Sun was to tilt them into a...

Thermochemical Analysis

A thermochemical analysis of the propellant exhaust products is necessary to determine the thermodynamic properties and composition. From the basic propellant formulations and chamber pressure, these parameters can readily be determined for equilibrium- or frozen-composition gas expansion by the method of Zeleznik and Gordon (ref. 82) or by the methods of references 83 through 86. These techniques are based on conservation of mass, Dalton's Law of Partial Pressures, adiabatic combustion, and an isentropic combustion process. The enthalpy, heat of formation, and free-energy data can be obtained from a current file of JANAF data (ref. 87). The species system usually is set to allow every gaseous species to be in the system of products selected from the thermodynamic file. Gaseous or liquid species are allowed to change phase at equilibrium temperature. Further discussion of the analysis is contained in reference 9.

Typical Beam Assembly

The hinged solar panels and stabilizing beamy are released from their restrained position by the ordnance system. Springs throughout the system provide the energy to deploy the solar panels. The deployment rate is controlled by a viscous danper in each wing section. The deployed position of the panels is maintained by positive latches. A rotary potentiometer indicates the position of the panels during deployment Development Tests 1 Solar Panel (SA-13)

Masterpieces of miniaturization

The SMART-1 spacecraft spans 14 metres with its solar panels extended, but otherwise everything for propulsion, communications, housekeeping and instrumentation fits into a cube just 1 metre across. Propulsion by an ion engine is not the only innovative technology on SMART-1. Its solar panels use an advanced type of gallium-arsenide solar cell in preference to the traditional silicon cells. And it will test new communications and navigational techniques. One set of solar panels is deployed for testing. Supports hold the panels from above to simulate zero-gravity. One set of solar panels is deployed for testing. Supports hold the panels from above to simulate zero-gravity. One cubic metre, 370 kg. Solar panels span 14 metres when deployed and provide 1.9 kW of power. 19 kg.

Further Protection From the

In addition to solar panels and propulsion system, many other components of the spacecraft needed protection from solar radiation. At the beginning of the program, considerable doubt existed that Kapton, a commonly employed heat-protection material, would survive in the anticipated environment. Kapton had been suggested as a replacement to the Teflon used on earlier Mariner spacecraft when engineering tests showed that Teflon failed at the intensities of solar radiation expected at Mercury. However, Kapton was found to become brittle with long exposure to temperatures above 354 C (670 F) and also in the environment of ultraviolet light and protons expected sunward of Earth's orbit.

The magic of ion engines

Operating in the near vacuum of space, ion engines shoot out a propellant gas much faster than the jet of a chemical rocket. They deliver about ten times as much thrust per kilo of propellant used. The ions that give the engines their name are charged atoms, accelerated by a choice of electric guns. If the power comes from the spacecraft's solar panels, the technique is called 'solar-electric propulsion'. Ion engines work their magic in a leisurely way. As solar panels of a normal size supply only a few kilowatts of power, a solar-powered ion engine cannot compete with the whoosh of a chemical rocket. But a typical chemical rocket burns for only a few minutes. An ion engine can go on pushing gently for months or even years - for as long as the Sun shines and the small supply of propellant lasts.

As on the International Space Station

In recent years, NASA has increased the frequency of extravehicular activity to prepare for the assembly of the International Space Station beginning in the late 1990s. Two recent extraordinary Shuttle missions have demonstrated the potential for EVA on the Space Station. These were the STS-61 and STS-82 missions in 1993 and 1997 in which astronauts conducted multiple spacewalks to service the Hubble Space Telescope. During each flight, crew members donned their spacesuits and conducted EVAs in which they replaced instruments and installed new solar panels.Their servicing activities required extreme precision and dexterity during EVAs lasting more than seven hours. As the Space Station evolves, astronauts will move components, such as antennas, to the most advantageous locations. For example, a communications antenna may become blocked when additional station modules are joined to the structure. Spacewalkers will have to traverse the outside of the modules and reposition the antenna...

The S truss is lowered into its transport canister at the Kennedy Space Center

Power and data cables and the thermal control system that provides heating and cooling wind through the 44-foot by 15-foot, 27,000-pound truss segment to carry energy and information to and from the station's extremities where solar panels collect electrical energy used to power experiments, computers, life support systems and other services. Video cameras, attached to the structure, monitor assembly operations and other activities on the station. Other instruments provide the data that astronauts and ground controllers use to maintain the station's position and orient the solar panels.

Concept Mission Profile

The solar electric stage delivers the customer payload to GEO without an AKM. Once delivered to LEO, the stage orients its solar arrays to maximize power production and successfully demonstrates its engine's ability to generate thrust. If a problem develops with the upper stage, the solar panels could be either retracted or ejected, and the upper stage could then be returned to Earth by the HRST. (Note the cost model did not account for such events, which were assumed to be rare.) The upper stage fires its solar electric engine and slowly shifts the orbit from LEO to GEO, using a low-thrust spiral-with-plane-change transfer maneuver. At low altitudes, its RCS subsystem may be used to compensate for atmospheric drag if Earth's upper atmosphere is expanded due to high solar activity. The one-way delta V is estimated at 19,030 ft s. For the 3,000-lb payload concept, the required burn time is estimated at 82.2 days (during 109.6 days) for the 10,000-lb concept, 79.2 days (during 105.6...

Interplanetary Cruise and Approach to Mars

Phoenix will begin the portion of its mission called the cruise phase after the spacecraft has established radio communications with Earth and sent information that the cruise solar panels are generating electricity and spacecraft temperatures are stable. This phase will last until three hours before Phoenix enters the atmosphere of Mars, more than nine months later.

Miscellaneous Usspacecom Support

USSPACECOM's SOI analysts determine the physical and dynamic characteristics of space vehicles by examining waveforms of signals collected by ground-based radar and electro-optical sensors. Items of interest to NASA include spin or tumble rates, rotation axis inertial orientation, stability mode (inertial or gravity gradient), changes from an earlier physical configuration (antenna or solar panel orientation), and reflectivity values, to name a few. Such information has been very valuable when collected on prospective rendezvous targets, such as WESTAR and PALAPA prior to STS-51A or Tethered Satellite during STS-75. Unfortunately, SOI data are generally classified, even when the objects tracked are not. If compromised, these data may reveal USSPACECOM's sensor and analysis capabilities. Therefore, requests for SOI data are only made when necessary, and USSPACECOM may or may not honor a request, depending on security issues or DoD priorities for use of SOI sensors. Since SOI data may...

Projected Design Changes

In fact, the reliability of the CC& S backup portion could be increased over that of the primary system, because the primary system would have more functions to perform, such as solar panel opening, etc. An overall reliability figure of 0.916 against a primary-function reliability of 0.71 would be preferred.

Capabilities Of Cassini

Saturn's gravitational field will be measured throughout the Cassini mission by radio tracking of the Orbiter trajectory at a variety of inclinations. The radius of Saturn at low and high latitudes will be determined by a series of radio occultation measurements. The far infrared spectrometer and microwave instruments will retrieve temperatures and their variation with latitude at the top of Saturn's convection zone. Combined with radio-occultation retrievals, the infrared measurements will provide a more accurate determination of the helium abundance. Thermal emission measurements by the composite infrared spectrometer(CIRS) and observations of reflected solar energy by the near infrared spectrometer will be used to refine our knowledge of the global energy balance.

Experiment Concept And Design

The temperature and the heat flux at the surface of the Moon are determined mainly by the solar energy impinging on the surface during one-half of the 29.5-day lunation cycle. During the lunar day, the surface temperature rises to approximately 380 K, which results in heat flow into the subsurface. After lunar sunset, the surface temperature drops to nearly 100 K, and heat flows out of the subsurface and is lost, by radiation into space. These very large temperature excursions, in part, are a result of the extremely low thermal conductivity and volumetric heat capacity of the fine rock powders that mantle most of the lunar surface (ref. 11-1) and, in part, are a result of the very tenuous atmosphere of the Moon. The low thermal conductivity of the bulk of the regolith (ref. 11-2) strongly inhibits the flow of energy into and out of the subsurface. At a depth of approximately 50 cm, the large surface variation of 280 K is attenuated to a nearly undetectable amplitude. At low lunar...

Cosmic Ray Experiment

A new set of cosmic ray detectors was carried to the surface of the Moon on the Apollo 17* mission. Two sets of detectors (including mica, quartz, glass, plastic, and foil) were exposed, one set facing the Sun and one set in the shade facing away from the Sun. During the time that the detectors were exposed, no significant solar activity occurred. Although the absolute flux levels for the 0.02- to 1-MeV amu energy range were considerably lower than those for the Apollo 16 mission, the shape of the spectrum is similar to that for the flare that occurred during the Apollo 16 mission and indicates that proportionate numbers of energetic particles are emitted by the Sun even during quiet periods. Heavy-element enrichment noted during flares is also present during the quiet periods. Tracks were also noted in the detectors facing away from the Sun. Because these particles also have a solar energy spectrum and presumably come from the Sun, the antisolar tracks indicate the existence of...

The P Integrated Truss Structure

Since the only readily available source of energy for spacecraft is sunlight, technologies were developed to efficiently convert solar energy to electrical power. One way to do this is by using large numbers of solar cells assembled into arrays to produce high power levels. The cells are made from purified crystal ingots of silicon that directly convert light to electricity through a process called photovoltaics.

Cloud Sat and Calipso Revealing the Secrets of Clouds and Aerosols

Together and separately, clouds and aerosols affect our climate in ways that are not completely understood. Clouds are a fundamental stage of Earth's water cycle, condensing water vapor and forming rain. They affect the amount of solar energy that reaches Earth, and the amount of energy that leaves Earth. They tend to cool Earth by reflecting sunlight back to space, while simultaneously warming the planet by absorbing and reemitting thermal radiation emitted by the surface and lower atmosphere. By modulating the distribution of heating within the atmosphere and at the surface, clouds fundamentally influence the circulations of the atmosphere and oceans. Shifts in this balance influence our climate. Aerosols affect where and how clouds form, and how much rain falls. Because we do not know enough about clouds and aerosols, there are uncertainties introduced into our predictions of climate change.

Station Reflections on Light and Heat

Scientists can understand more about the surface composition and texture of asteroids by measuring how much visible and infrared energy is reflected or emitted. Scientists can measure the surface temperatures of asteroids by measuring how much absorbed solar energy they emit as infrared energy.

Nasa Plans To Put An Aura Around The Earth

As the composition of Earth's atmosphere changes, so does its ability to absorb, reflect and retain solar energy. Greenhouse gases, including water vapor, trap heat in the atmosphere. Airborne aerosols from human and natural sources absorb or reflect solar energy based on color, shape, size, and substance. The impact of aerosols, tropospheric ozone and upper tropospheric water vapor on Earth's climate remains largely unquantified. Aura's ability to monitor these agents will help unravel some of their mystery.

Mars Surface Operations

If electrical output from the solar panels remains adequate and other subsystems are functioning, the mission for Phoenix might be extended for an additional month or two, into late summer or early fall at the landing site. Factors in how long the lander can keep getting adequate solar power include whether it lands with a southward tilt and how quickly dust accumulates on the solar panels. However, Phoenix will not be capable of outliving its prime mission several times over, as other recent Mars missions have done. Mission planners anticipate that by about Sol 150, a combination of less sunlight per day and accumulated dust on the solar panels will shrink electrical output below the amount needed for heating to keep Phoenix operating. Within a few more months, carbon dioxide frost will heavily coat the region of the landing site and the spacecraft itself.

Site Selection Consideration

The idea that water ice may exist at the poles of the Moon has gained currency in the last decade as a result of two robotic missions (Clementine and Lunar Prospector) that found evidence for enhanced volatile concentrations associated with the poles. If such water ice exists, the extraction of water from this lunar resource requires at least two orders of magnitude less energy than does the synthesis of water from the hydrogen autoreduction of solar wind gas-saturated regolith. Thus, the polar deposits qualify as high-grade ore. Moreover, there is evidence that certain small areas near the poles may be in near-constant sunlight, providing sites that have access to continuous solar power and are also thermally benign.

Mounting ring for the spacecraft installation on the launchvehicle

On the outside of the assembly section there are a radiator of the thermal control system, 4 approach and orientation engines with a thrust of 10 kg each, 8 orientation engines with a thrust of 1 kg each, and rear attachment points of the solar panels. - onboard color orientation lights (on the ends of solar panels).

Passive Seismic Experiment Package

The PSEP also includes the solar panels, which power the experiment during the lunar day and isotope heaters which will help it survive the cold (-300 degree F) lunar night. The solar panels convert the energy of sunlight to electricity and have an output of 33 to 43 watts. The panels operate only durinc the lunar day during the lunar night, PSEP is inoperative. The output of the solar panels is fed to the power conditioning unit and supplies all the electrical requirements of the PSEP excess current is dissipated by

System Configuration And Assembly

Major elements of the VO were the bus structure to which all the other elements attach, the VLCA structure, the truss adapter which attached the V S C to the Centaur LV, and VO appendages such as the four two-piece solar panels, the scan platform with the VO science instruments, the HGA, the LGA, and the bolt-on propulsion module. A schematic diagram showing the mechanical configuration is shown in figure 1. The major subsystems, along with the abbreviations and locations on VO, are given in table 1. The electronic packaging arrangement

Space Environment Monitor

The SEM-2 provides measurements to determine the intensity of the Earth's radiation belts and the flux of charged particles at satellite altitude. It provides knowledge of solar terrestrial phenomena as well as warnings of solar wind occurrences that may impair long-range communications and high-altitude operations, damage satellite circuits and solar panels, or cause changes in drag and magnetic torque on satellites.

Radiation Environment

Important data concerning space radiation and its effects on materials, particularly silicon diodes and glass, have been obtained from the flight experience of solar panels. The development of radiation resistant solar cells and cover glass has also benefited sun sensor development because comparable parts are used. Other radiation effects that may be significant in sun sensor design are the embrittlement of wire insulation, the discoloration of painted surfaces, and the bleaching and flaking of black anodized surfaces.

Spacecraft Integration

The location of each damper component should be determined by considering the type of damper and the spacecraft configuration. In single part dampers, the damper rod should be attached either to the main spacecraft structure or to the solar panels. In two-part dampers, all components that require electrical power and those requiring magnetic balancing with respect to the spacecraft should be located on that part of the damper that is fastened to the spacecraft. For example, the photocells and the light source of the relative angle sensor and the magnets of the damping element should be attached to the spacecraft, while the mating components should be attached to the damper boom.

Viking Orbiter Launch Operations Results

Following completion of the system test phase, VO-2 was built up as much as possible before being moved to ESA-60 (see Table 7). At ESA-60, the PROP, solar panels, and thermal blankets were installed. Various special tests, a precountdown test, and a biosampling were performed before the completed orbiter was moved to SAEB-2 for lander mating and encapsulation. A number of problems occurred during the final assembly phase, some of them involving the facilities. These problems are discussed below.

Solar Array Subsystems

Mechanism assembly a cassette drum to hold solar panels, a cushion to protect the blanket, and motors and subassemblies. The assembly rolls out the blanket, applies tension evenly so the blankets stretch, and transfers data and power along the wing assembly. The blanket can roll out completely or part way. The secondary deployment mechanism also has a manual override (see Fig. 5-29).

Voyager Encounter Of Neptune

Two key spacecraft characteristics made the extra planetary encounters possible reprogrammable onboard computers' receptive creative software engineering and three radioisotope thermoelectric generators -- devices that convert the heat from the radioactive decay of plutonium 238 into electricity to power the spacecraft components and instruments. The radioisotope thermoelectric generators allow the Voyager spacecraft to operate in regions of the solar system where solar panels cannot be used.

Chapter Spacecraft Classification

Spacecraft Power Generation

A spacecraft designed to travel to a distant planet and enter into orbit must carry with it a substantial propulsive capability to decelerate it at the right moment to achieve orbit insertion. It has to be designed to live with the fact that solar occultations will occur wherein the planet shadows the spacecraft, cutting off solar panels' production of electrical power, and subjecting the vehicle to extreme thermal variation. Earth occultations will also occur, cutting off uplink and downlink communications with Earth. Orbiter spacecraft are being used in the second phase of solar system exploration, following up the initial reconnaissance with in-depth study of the planets. These include Magellan, Galileo, Mars Global Surveyor, and Cassini.

And Rendezvous And Dock

Docking Mechanism Apollo

The AR& D scenario for rendezvous and docking the Soyuz or the Progress vehicles to the space station Mir is as follows.20 The process begins with the Mir transmitting a beacon radio frequency (RF) signal from hemispherical-coverage antennas on the ends of its solar panels. The chase vehicle, which could be either the Soyuz or the Progress, has a gimballed, 0.5-meter dish antenna that searches for this signal. The 0.5-meter dish antenna system can detect and acquire it up to 200 kilometers away. Once this is accomplished, the gimballed antenna then begins to angle track the signal from the Mir. At this point, the RF beacon signal is turned off and a transponder on the Mir is connected to the antennas on its solar arrays. The chase vehicle now uses the return signal from the transponder to determine range using the time delay and range rate using the Doppler shift of the returned signal. Using this information, the chase vehicle closes in on the Mir until a range of about 200...

Thrust Vector Control

A problem common to all spacecraft using TVC is that of transients caused by engine ignition and thrust tailoff which can result in oscillatory response of the attitude-control system. Difficulties of this nature were encountered in the design of the Mariner Mars spacecraft which used TVC provided by thrust deflection. The Mariners 4, 5, 6, and 7 spacecraft basically were modified Ranger spacecraft. The primary changes were modification of the engine to provide the capability for two midcourse burns (ref. 31) and relocation of the engine from the principal spacecraft axis (roll) to a location between the pitch and yaw axes (fig. 12). Analysis of the Mariner autopilot included consideration of the structural dynamics of the solar panels and a scan platform. The results of a computer simulation indicated that the structural dynamics of the scan platform when free to rotate could cause autopilot transient response during the ignition and thrust tail-off portions of the midcourse burn....

Command and Data Handling

Genesis Spacecraft

Solar panels Solar panels All of the spacecraft's power is generated, stored and distributed by the electrical power subsystem. The system obtains its power from an array of standard silicon solar cells arranged on two panels on either side of the equipment deck. The two solar panel

Power Capabilities And Requirements

The VO had 2.4-kHz single-phase, 400-Hz three-phase, regulated dc (30 V and 56 V) and unregulated dc (25 V to 50 V) power sources. Unregulated dc power was also provided for the VLC. Arrays of photovoltaic cells arranged on four double-section, folding solar panels furnished primary power for all Sun-oriented operations. Two identical nickel-cadmium batteries were used as a secondary source of power for off-Sun operations and to share the load when power demand exceeded the solar array capability. Redundant power conditioning and distribution functions were provided with two battery chargers, two booster regulators, two 2.4-kHz inverters, two 400-Hz three-phase inverters, two 30-V dc converters, and associated power source logic and control and switching functions. (See the simplified block diagram in fig. 6.) The hardware, operating modes, and performance are described in detail in the section Power Subsystem. The VO unregulated (raw) power bus was supplied by solar panels and...

Description Of Major Subsystems

Structure subsystem STRUS provided mechanical support and alignment for all flight equipment, passive thermal control, and micrometeoroid protection. In addition, the structure provided means for handling the assembled VO for flight qualification testing, transporting, and mating operations with the VLC and the LV. STRUS included a basic bus structure of irregular octagonal shape about 2.4 m (8 ft) maximum across the flats substructures for the PROPS, the solar panels and outriggers, the scan platform, the electronic assembly chassis, the high- and low-gain antennas (except feed assemblies), and the V S C A and VLCA and the micrometeoroid and thermal blankets. (See fig. 1.) Power subsystem PWRS provided the VO with 2.4-kHz single-phase, 400-Hz three-phase, regulated dc (30 V and 56 V), and unregulated dc power. Unregulated dc power was also provided for the VLC prior to separation. PWRS utilized arrays of photovoltaic cells arranged on four double-section, folding solar panels to...

The International Space Station

More than four times as large as the Russian Mir space station, the completed International Space Station will have a mass of about 1,040,000 pounds. It will measure 356 feet across and 290 feet long, with almost an acre of solar panels to provide electrical power to six state-of-the-art laboratories.

Lunar Orbiter Program

Lunar Orbiters III and V electrical power systems, like their predecessor Lunar Orbiter I, were similar In configuration 4,5 except for the addition of the booster regulator for the photo subsystem in Lunar Orbiter V spacecraft. See Figure 11-2. The solar array for both Lunar Orbiters III and V functioned normally during the extended mission, providing sufficient power to maintain a constant bus voltage of 30.56 volts dc when the solar panels were directed toward the sunlight. The total solar panel power for Lunar Orbiter III and V at launch was 13.30 and 12.49 amperes at 30.56 volts, respectively. The battery performance was as predicted throughout the mission.

Entry descent and landing

About 10 minutes before entry, the spacecraft will be commanded to switch to inertial navigation - computing its position, course and speed from gyroscopes and accelerometers. Six minutes before entry, the spacecraft will fire its thrusters for 80 seconds to turn it to its entry orientation. Five minutes before entry and 10 minutes before landing, the cruise stage will separate from the aeroshell-encased lander. Cut off from the cruise stage's solar panels, the lander will rely on its internal battery until it can unfold its own solar panels on the planet's surface. The Deep Space 2 microprobes, piggybacking on the lander's cruise stage, will be jettisoned about 18 seconds later. The lander will then be commanded to assume the correct orientation for atmospheric entry. The pulse-modulated descent engines will maintain the spacecraft's orientation as it descends. The engines will fire to roll the lander to its proper orientation so that it lands with the solar panels in the best...

International Space Station General Resource Reel July Reference Master JSC Cut sheet available

This comprehensive reel includes all major recent station video, including new animation of the complete station, actual footage of the Zarya (first element) launch, STS-88 Highlights of the first station assembly mission to join the Zarya and Unity modules in space, STS-96 and STS-101 mission highlights, Zvezda service module, STS-92 Animation of Z-1 and PMA installation, Z-1 truss video, STS-97 animation of P6 Solar array installation and deploy, Solar panel at KSC, Long Spacer at KSC, new Expedition One crew training with Orlan and American suits training underwater at JSC, the Expedition One crew training in Russia, U.S. Lab footage at KSC and Marshall, Multi Purpose Logistics Module at KSC, new footage of the Canadian robot arm, the U.S. Airlock , ISS Solar array, Japanese Experiment module, Animation of the Columbus Attached Pressurized Module, new TransHab animation and real video of testing at JSC, U.S. Hab video from Marshall, X-38 free flight test from March 5, 1999, and...

Test Test and Test Again

A most critical item to the success of the mission was the solar panel system, since if these panels did not move from the stowed to the operating position, the spacecraft would be starved of electrical energy and could not operate. Viscous boost dampers had to be developed for the solar panels as well as for the magnetometer boom to prevent the two systems from banging together during the launch. The objectives of this development program were to verify that the damping force would meet the requirements for the spacecraft over the operational range of vibration frequencies, and to develop assembly techniques for the magnetometer boom damper. During the development program, both dampers were subjected to small- and large-amplitude vibration testing at many frequencies. These tests were successful and showed that the dampers would safeguard the solar panels and the magnetometer boom. A development test fixture was used to verify that the release mechanism exerted sufficient force to...

Concept Selection Criteria

This system uses a deployable concentrator to convert solar energy to thermal energy. The solar energy is focused into the engine absorber where propellant (H2) is heated to provide thrust to the stage. The solar thermal upper stage (STUS) requires multiple burns (300 to 1,000) of the engine, and on-orbit cryogenic fluid management to complete its mission, which may be lengthy. This system uses a deployable concentrator to convert solar energy to thermal energy. The solar energy is focused into the engine absorber where propellant (H2) is heated to provide thrust to the stage. The STUS requires multiple burns (300 to 1,000) of the engine, and on-orbit cryogenic fluid management to complete its mission, which may be lengthy. This reusable system could get propellant (LH2) resupply either on-orbit or on ground.

Solar Thermal Propulsion

Solar thermal propulsion is an advanced propulsion concept that utilizes concentrated solar energy to heat a propellant and thereby produce thrust. This propulsion concept will be used on an upper-stage vehicle with the main application being the transfer of payloads from low-Earth orbit to geostationary Earth orbit. The solar thermal propulsion concept is characterized as having low thrust (2 to 5 pounds force), high specific impulse (800 to 1,000 seconds), and is a much more efficient means of transporting payloads when compared to conventional chemical propulsion systems. Such efficiency will allow for the reduction of launch costs and for the increased competitiveness of the U.S. launch services industry.

Data Subsystem Details

Physically, the DSS is contained in the pallet on which the PSE and solar panels are mounted. Each component is boxed, and is interconnected by a wire harness ending in multi-pin connectors which make removal and replacement of individual components a relatively simple operation.

Mariner Reliability in Flight

The short which occurred in the solar panel on October 31 may have been the result of poor assembly techniques or inadequate design. No judgment can be stated here as to the cause of the short however, this may be a good example of the degradation of system reliability due to some cause other than a mere electronic-component failure. While this condition existed, the transponder output fluctuated, the spacecraft magnetic field changed, and the power system was forced to work under abnormal conditions. The actual seriousness of each of these anomalies is difficult to assess. Had the condition occurred much earlier in flight, the spacecraft power system might have gone into a load-sharing mode, from which recovery would have been doubtful. However, it should be noted that the panel design accommodated a power pad to provide for radiation degradation and cell damage. Solar panel short circuits

Handling and Contamination

A sensor is not out of danger from rough handling and contamination after it is installed on the spacecraft. For example, the sun sensors of the Radio Astronomy Explorer were mounted on their respective solar panels before the start of the panel test program. Hence, the sensors were subjected to needless handling, and four out of the eight sensors were damaged because the test personnel were not made fully aware of the sensor's susceptibility to damage. Illumination interference on Nimbus 1 caused the spacecraft's solar panels to deviate by about 20 from their proper orientation with respect to the Sun. This problem was traced to a thermal shield that extended further into a sensor's field of view than had been anticipated because the planned circular orbit was not obtained. When the spacecraft attitude deviated by 1 or 2 from local vertical, the shield shadowed the sensor supplying orientation signals for the panels. The Sun shield was redesigned for later spacecraft so that attitude...

The Solar Arrays Unfurl to their Full Length

There are two solar array wings on the P6 module, each deployed in the opposite direction from each other. Each SAW is made up of two solar panels mounted to a common mast. Prior to deployment, each panel is folded into a Solar Array Blanket Box (SABB) measuring 20 inches high and 15 feet in length. The SSU is designed to coarsely regulate the solar power collected during periods of insolation - when the arrays collect power during sun-pointing periods . A sequence of 82 separate strings, or power lines, leads from the solar array to the SSU. Shunting, or controlling, the output of each string regulates the amount of power transferred. The regulated voltage setpoint is controlled by a computer located on the IEA and is normally set to around 140 volts. The SSU has an overvoltage protection feature to maintain the output voltage below 200 V DC maximum for all operating conditions. This power is then passed through the BMRRM to the DCSU located in the IEA. The SSU measures 32 by 20 by...

Mars Polar Lander spacecraft

The lander obtains its power from a total of six solar panels. The four larger panels are arranged as a pair of wings on either side of the lander, and are deployed after landing. Two smaller panels fixed to the side of the lander were added to increase the total power output after the main solar panels were made as large as possible while still fitting in the launch vehicle's fairing or nose cone. During the southern Martian summer at the time of arrival, the sun never sets below the horizon at the landing site. A rechargeable 16-amp-hour nickel-hydrogen battery will keep the central electronics enclosure relatively warm (above -30 C (-22 F)) during -80 C (-110 F) night-time temperatures near the Martian pole. The lifetime of the battery will probably be the main factor determining how long the lander operates. As nights grow colder in late Martian summer, the battery will eventually be unable to provide enough power to keep the spacecraft enclosure warm at night. The lander...

Temperature Control

Solar panel Figure 32 shows the solar panel, associated components (deployment damper, rate limiter, etc.), and temperature sensor locations. The back side of the solar panel was painted white. The deployment dampers were blanketed and the portion not covered by the multilayer blanket was polished. The rate limiter, release rod, and latch assembly had a polished finish. Two of the horizontal members (normal to the Sun) of the outriggers had a blanket shade. Figure 32.- Solar panel. Figure 32.- Solar panel. Attitude control jet and acquisition Sun sensor A pattern of white paint and polish characterized the attitude control jets and accompanying ACQ SS. Fiberglass standoffs were used to isolate the package from the solar panel. Each manifold was heated during occultation. The Sun sensor was coupled to the jets by four polished aluminum standoffs. No temperature sensors were provided for these appendage items. Sun gate and cruise Sun sensor The Sun gate and CR SS were mounted together...

By Sun power to the Moon

It's not very big, just a ' box a metri W with folded solar panels attached. Six strong men could lift it. It weighs less than 370 kilos, compared with thousands of kilos for Ariane's usual satHitfljSo it should pose no problems as an auxiliary passenger. Its main purpose is to let engineers evaluate a new way of propelling spacecraft anter-fanging space missions. Power from SMART-1's solar panels will drive an electric propulsion system called an 'ion engine'. The demonstration task is to overcome the Earth's gravity and put the spacecraft into orbit around the Moon. Solar panels of the gallium-arsenide type that will power SMART-1 enabled the Dutch-built solar-powered car 'Nuna' to win the World Solar Challenge race across Australia in 2001. esa Solar panels of the gallium-arsenide type that will power SMART-1 enabled the Dutch-built solar-powered car 'Nuna' to win the World Solar Challenge race across Australia in 2001. esa

The Voyager Mission

Voyager Trajectory

Because they would be traveling too far from the sun to use solar panels, the Voyagers would use radioisotope thermoelectric generators. These devices, used on other deep space missions, convert the heat, produced from the natural radioactive decay of plutonium, into electricity to power the spacecraft instruments, computers and radio.

Plans For Future Improvements In Marsgram

Since the current Mars-GRAM relies on estimates of the daily total surface solar irradiance for the estimation of daily maximum, average and minimum temperatures, it would be a straightforward addition to Mars-GRAM to add a subroutine which calculates the direct (beam) and diffuse (scattered) components of the time-dependent solar irradiance at the surface. Such a routine for estimation on the surface solar irradiance would be based on the model approach of Justus and Paris (1985), with parameters adapted for conditions of the Mars atmosphere and dust optical properties. The influence on solar irradiance due to dust storm conditions could easily be incorporated, since optical depth has been measured at the Viking Lander sites (Tillman et al., 1979 Zurek, 1982), and other dust optical properties have also been inferred (Toon et al., 1977). Addition of a surface solar irradiance module for Mars-GRAM would be very useful for such applications as (1) analysis of thermal heating...

The Spacecraft and their Paths

Mcdonnell Spacecraft Sequential Diagram

The Viking orbiter is also equipped with a larger area of solar energy collecting cells to provide additional electrical power for the complex mission a total of 620 watts. A large parabolic antenna is motor driven to point toward Earth and provide a tight communications beam that allows rapid transmission of data to the big antennas of the Deep Space Net located in California, Spain, and Australia. Solar Energy Controller Solar Panel

Pioneer Venus Mission

The solar panel main and charge array characteristics are shown in Table 8.4 2 along with other Orbiter spacecraft and Multiprobe Bus power The Orbiter bus power subsystems and thermal subsystem were interfaced with each other in order to stabilize spacecraft unit temperature by dissipating excess solar panel power. The Orbiter and Multiprobe main solar arrays were designed to supply electrical power to the subsystem within the voltage range of 28 volts dc + 10 under varying conditions of sun angle, temperature, and solar intensity. Each solar panel had two smaller battery charge arrays that were connected in series with the main array to supply 36 volts to the battery chargers. The probe power subsystem of the Multiprobe system was a silver-zinc battery that was located within each probe. Before the probes were separated from the bus, they received power from the bus spacecraft solar panel and or nickel-cadmium batteries. Once the power was switched from the bus to the probes, a 40...

Test Battery Removed To Install On Vo Reinstalled

The orbiter cable installation basically consisted of a solar energy controller (SEC) harness on the aft end of the orbiter, an upper ring harness, and interconnecting cables to each bay. The electronics assemblies were installed in as logical a sequence as possible as they became available, either as breadboards or prototypes. Pyrotechnic equipment, scan platform, science in strum ents, guidance sensors, and mechanical devices were installed and individually exercised as described below in preparation for system testing. Not installed at this time were solar panels, radio antennas, propulsion subsystem, and reaction control assemblies.

This line art highlights the S truss and Mobile Transporter additions that will be made to the International Space

Preparation Pyrotechnic Device

Atlantis will carry the first major external truss section for the station, a 43-foot-long girderlike segment that will lay the foundation for an eventual cross-beam that will stretch more than 350 feet. Nine additional truss segments will be linked on future missions to the centerpiece segment carried by Atlantis to form the finished structure. The finished truss will support almost an acre of solar panels and giant cooling radiators. Although the International Space Station already is a fully functional research complex with a single United States laboratory, the additional solar panels and radiators will provide the electricity and cooling necessary for Japanese and European laboratories to be attached to the station as well as a future U.S. centrifuge laboratory.

In situ calibration and validation

Once Beagle 2 has finally come to rest on the surface of Mars and fully deployed itself, the PAW is ready to obtain the first image of the landing site. This is achieved via the Wide Angle Mirror, which moves into the FOV of the right-hand stereo camera when the pre-tensioned spring holding it down is released by the opening of the lid and solar panels. The figure of the WAM is designed to provide a 360 view of the landing site that includes the horizon. Fig. 18 shows simulated and actual views using the WAM. A main hinge mechanism allows the shells to open and provides self-righting by shifting the centre of gravity (dominated by the base) if the lander lands the wrong way up. The lander lid houses the solar panels, which are deployed via electrically-powered hinges. The base shell accommodates the robotic arm, which allows the PAW instruments to be deployed. Beagle 2 is powered via a 42-cell lithium ion battery that is kept warm by the thermal insulation and by heater power during...

Spacecraftmission Mir PE

To save weight, the Mir base block was launched with only two solar arrays. These provided a total of only 9.4kW of electricity, leaving the base block hungry for power. The base block had a socket on top for a third array delivered inside the Kvant module. On June 9 Romanenko and Laveikin capped a busy day spent studying Supernova 1987A using the Kvant module's Roentgen observatory by undergoing a medical checkup to certify them fit for two solar array installation EVAs. Spacewalk preparations alternated with astronomical observations on June 10 and 11, and lasted all day on this date. The first installation EVA began late in the evening. Romanenko and Laveikin attached an extendible hinged lattice girder truss to the top of the Mir complex, then attached folded solar panels to both sides of the girder. To test their ability to operate without EVA restraints, Romanenko and Laveikin employed no foot restraints on this and their next EVA, relying instead on tethers. Laveikin later...

Surface Temperatures Deduced From Thermocouple Measurements

The first term on the right side of equation (9-13) represents flux into the cable element from the lunar surface the second term represents direct flux from the Sun the third term represents solar energy ieflected diffusely from the lunar surface and impinging on the cable.

Return to the Innermost Planet

Attitude control gas, and analysis showed that gas usage would have to be reduced well below the normal cruise rate to last until Mercury III. Further, multiple trajectory correction maneuvers had to be conducted (five were finally required between the first and third encounters), meaning that a way had to be found to use the gyros without causing the oscillation problem. The only technique available to reduce gas usage, which involved using the solar panels and high-gain antenna as solar sails, was as yet little understood. Only limited experience had been gained so far during the mission with regard to this unplanned method. The prospects of achieving a third encounter therefore seemed dim, and Mercury II, although far more likely, was by no means assured. After Mercury II, Mariner 10 was placed back in the cruise mode in which the high-gain antenna and solar panels were used as light-pressure torquers to save attitude control gas for a third encounter. The antenna was placed in a...

Active Closed Loop Control

Spacecraft Attitude Control

Eliminates skewed high thrust control axes due to thrustor location on solar panel. 3. Eliminates reaction control coupling with solar panel flexibility. a. Performance best with thrustors on the tip of solar panels. b. Solar panel temperatures extreme. 4. Introduces thrust vector control (TVC) coupling with solar panel flexibility.

New Horizons Nuclear Safety

RTGs enable spacecraft to operate at significant distances from the Sun or in other areas where solar power systems would not be feasible. They remain unmatched for power output, reliability and durability by any other power source for missions to the outer solar system.

Spectral Content of Albedo and Long Wave Radiation

Spectral Albedo

Frequently cited evaluations of the spectral distribution of solar radiation reflected by the Earth-atmosphere system are those of Fritz (ref. 15), Hewson (ref. 32), and Coulson (ref. 33). Figures 1, 2, and 3 show results obtained in 1969 by W, Hovis and M. Forman of the NASA Goddard Space Flight Center with a spectrometer aboard the NASA Convair 990 aircraft. The figures show typical values of relative reflectances* associated with various cloud formations, unclouded ocean regions, and wheat fields under a clear sky. Figure 1 shows the variations in reflection of solar energy among several types of clouds. It can be seen that the thickest, highest cloud reflects the most energy, and the snowing cloud the least. The peak value of reflected solar energy occurs at about 0.58 xm, Figure 2 illustrates the variations in reflectance spectra from two very similar areas (wheat fields) when the measurements were taken only a minute apart. The spectra given in figure 3 for ocean areas under...

Other Magnetospheric Emissions

We have seen how energetic particles from the solar wind are one source of energy. Both nonuniform and nonthermal plasma distributions represent additional sources of free energy. Generally, waves that grow at the expense of a nonthermal or nonuniform feature interact back on the plasma distribution to try to eliminate the nonuniform or

Viking Mission Plan

Two identical Viking spacecraft, each consisting of an Orbiter and a Lander, were launched from Kennedy Space Center on a Titan III, Centaur from Launch Complex 41. Viking 1 (Orbiter 1 Lander 1) was launched August 20, 1975 and Viking 2 (Orbiter 2 Lander 2) was launched September 9, 1975. With a second firing of the Centaur engine the Vikmg spacecraft were injected into trans-Mars trajectories. Just after the cruise to Mars began, the spacecraft separated from the Centaur, and the Orbiters then deployed antennas and solar panels and acquired the sun and the bright star Canopus for navigation. During cruise, the Orbiters were the primary operating portions of the spacecraft, supplying power to the Landers, maintaining proper attitude for communication to Earth and thermal balance, commanding Lander checkouts and operations, and relaying Lander data to Earth for analysis of Lander hardware status. The Orbiters also performed midcourse maneuvers to insure that the spacecraft arrived at...

Objective Eight Chart our destiny in the Solar System

Which will observe the Sun's outer layers to determine its interior dynamics and the activity of the solar corona, the source of sunspots and active regions, and origin of coronal mass ejections. A second LWS component is a constellation of Sentinels around the Sun to observe the movement and evolution of eruptions and flares from the dynamic Sun through the interplanetary medium to Earth's orbit. LWS geospace missions are the Radiation Belt Mappers and the Ionospheric Mappers. The Radiation Belt Mappers will characterize the origin and dynamics of terrestrial radiation belts and determine the evolution of penetrating radiation during magnetic storms. The LWS Ionospheric Mappers will gather knowledge of how Earth's ionosphere behaves as a system, linking incident solar energy with the top of Earth's atmosphere.

Thin Filmlhrge Phvoff

In the case for space, ultra-thin film can be tightly packaged for launch, then deployed in orbit to establish large solar concentrators and antennas for geographical surveys and communications. Lightweight, inflatable structures will soon act as huge reflectors that focus the sun's energy to heat propel-lant for thrusting payloads to high altitudes. Gossamerlike thin film might take the shape of collecting and transmitting dishes for 21st century satellites that collect solar energy, then beam the energy to Earth.

Section Configuration

Thus, economic factors, including the use of previously developed Apollo and Gemini hardware and the retention of hardware started for the wet workshop, had a major influence on the conceptual development aud final design of the Saturn Workshop. When launched, it contained all the elements needed to sustain the crew and operations. Food and water arp stored on board and the necessary hardware if provided for collection and disposal of human wastes. There are provisions for supplying a breathable atmosphere and controlling temperature, pressure, and humidity. Electrical power is produced by direct conversion of solar energy, using two sets of solar arrays, one mounted on the workshop and the other on the solar observatory. The necessary system sensors, controls, c1 communications devices are included so that decisions can be made and implemented either by the crew or ground controllers. Means of controlling precisely the orientation of tl- Saturn Workshop are provide ., to facilitate...

Hubbles New Solar Arrays

The high efficiency solar panels have supporting frames made of aluminum-lithium, which is stronger and lighter than the type of aluminum commonly used in spacecraft construction. The supports are less sensitive to the extreme temperature changes of Hubble's harsh environment. During each 97-minute orbit, Hubble spends about two-thirds of its time in searing sunlight and the other third in the frigid darkness of Earth's shadow. The rapidly cycling conditions cause the temperature of the solar panels to fluctuate between minus 94 degrees Fahrenheit (minus 70 degrees Celsius) and 187 degrees Fahrenheit (86 degrees Celsius). The solar arrays reach their hottest temperature just 10 minutes after leaving the chill of Earth's shadow. The Hubble program bought the new solar panels from the production line of a commercial system of communications satellites. At NASA's Goddard Space Flight Center in Greenbelt, Md., four of these panels were attached to an aluminum-lithium support wing...

Launch and Acquisition Phase

The Delta's second stage engine will cut off for the first time at first second-stage engine cutoff will occur at 7 minutes, 16.5 seconds after launch. After coasting, the second stage will restart at launch plus 38 minutes, 35 seconds, and will burn for 275 seconds. At this point, the observatory and the second stage will have escaped Earth's gravity and will be in the proper Earth-trailing orbit. Five minutes later, the observatory will separate from the second stage. The second stage will perform a maneuver to move away from the observatory to avoid future contact and possible contamination. After separation from the second stage, the observatory will orient itself so the Sun shines on the solar panel assembly for generating power. Low-gain antennas will be used for communications during at least the first 24 days of operations because the angle between the observatory, Earth and the Sun will not allow use of a dish-shaped high-gain antenna. The telescope must be shielded from the...

Experiment Equipment

A-4) consisted of three long-period seismometers and one short-period vertical seismometer for measuring meteoroid impacts and moonquakes and to gather information on the moon's interior for example, to investigate for the existence of a core and mantle. The passive seismic experiment package had four basic subsystems the structure thermal subsystem to provide shock, vibration, and thermal protection the electrical power subsystem to generate 34 to 46 watts by solar panel array the data subsystem to receive and decode Network uplink commands and downlink experiment data and to handle power switching tasks and the passive seismic experiment subsystem to measure lunar seismic activity with long-period and short-period seismometers which could detect inertial mass displacement. Also included in the package were 15-watt radioisotope heaters to maintain the electronic package at a minimum of 60 F during the lunar night. A solar panel array of 2520 solar...

Chemical Weathering

On Mars, the alteration of material in impact ejecta blankets or volcanic geothermal regions would also have a distinctive geochemical signature due to the circulation of hydrothermal fluids interacting with impact glasses and breccias. Calculations based on minimizing the Gibbs free energy of the chemical system by Zolensky et al. (1988) show that gibbsite (Al(OH) 3), kaolinite (an Al-rich clay), and nontronite (an Fe-rich smectite clay) would be present. The abundance of each clay would depend on the amount of rock that has reacted and the initial composition of the rock. If only a small percentage of the rock reacts, carbonates would not form.

Electrical Power

The solar panel array provides the electrical power needed to operate the observatory for five years. The array consists of two solar panels, each with 392 solar cells. Each Solar panel Solar panel shiel Solar panel Solar panel shiel solar cell is 5.5 by 6.5 centimeters (2.2 by 2.6 inches). Together the cells can convert radiation from the Sun into a total of 427 watts of electrical power at the beginning of the mission and 386 watts towards the end. Unlike most spacecraft solar arrays that are deployed shortly after launch, this mission's solar array is fixed. To ensure that sunlight will hit the solar panels properly, the telescope cannot be pointed at targets more than 120 degrees away from the Sun. The solar array also shades the telescope from direct exposure to the Sun. Half of the solar array's surface area is covered with solar cells. The other half is covered with flexible optical solar reflectors that reduce the overall solar panel temperature to about 57 C (134 F).

Ranger Project

Was used, and thus no provision was provided for recharging the batteries when the solar panels were generating power. The solar panels supplied power over the entire mission after Sun acquisition except during the midcourse maneuver phase. Each solar panel consisted of 4896 silicon solar cells. During the midcourse maneuver, the solar panels supplied the raw power load of approximately 145 watts out to a pitch ingle of 48 degrees. The electrical load was shared between the solar panel and battery during this period. After the midcourse motor burn and Sun acquisition, the solar panels supplied the total raw power load of approximately 120 watts at a pitch angle of 58 degrees 3 .

Overview

The observatory is composed of three main sections. The tube-shaped cryogenic telescope assembly includes the telescope and scientific instruments, enclosed within a protective outer shell. At the rear of the telescope is an eight-sided spacecraft structure that houses all the computers, electronics, antennas and thrusters needed to keep the observatory operating and oriented correctly in space. The spacecraft's third main section is the solar panel array, which serves as the observatory's power plant and The outer shell that encloses the telescope serves as both a dust cover and a heat shield. Shaped like a cylinder, the entire outer shell is composed of aluminum - a quarter inch layer in a honeycomb pattern is sandwiched between two sheets. The side of the outer shell that faces the Sun has a shiny silver coating to reflect away heat from the solar panels. The side facing away from the Sun has a black coating designed to radiate any residual heat from the solar panel and spacecraft....

Heatshield

Features The lander will parachute to the surface. One second before impact on the Martian surface, three airbags will inflate on each of the three folded petals of the lander, cushioning its impact. After the airbags have deflated, the petals then deploy, exposing solar panels to the sunlight, and righting the lander. The rover is then deployed by driving off the solar panel and onto the Martian soil. The lander is designed to operate on the surface for over 30 Martian days and nights, and return a panoramic view of the Martian landscape. It will also measure the soil's chemistry and characterize the seismic environment.

Viking Orbiter

The combined weight of the Orbiter and Lander was one factor that contributed to an 11-month transit time to Mars, instead of five months for Mariner missions. The longer flight time then dictated an increased design life for the spacecraft, larger solar panels to allow for longer degradation from solar radiation and additional attitude control gas. The Orbiter is 3.3 m (10.8 ft.) high and 9.7 m (32 ft.) across the extended solar panels. Its fueled weight is 2,325 kg (5,125 lbs.). The Orbiter is stabilized in flight by locking onto the Sun for pitch and yaw reference and onto the star Canopus for roll reference. The attitude control subsystem (ACS) keeps this attitude with nitrogen gas jets located at the solar panel tips. The jets fire to correct any drift. A cruise Sun sensor and the Canopus sensor provide error signals. Before S*un acquisition four acquisition Sun sensors are used and then turned off. Relay from the Lander. through an antenna mounted on the outer edge of a solar...

Spacecraft Structure

The solar array structure consists of the yoke, two solar panel substrates, and the trim tab panel. The yoke structure is made of GFRP beams. The solar panel substrates are of sandwich construction with lightweight aluminum honeycomb core and thin GFRP faceskins. A thin Kapton film (3-mil thick) is bonded to the solar cell side of the substrates to electrically insulate the solar cells from the graphite faceskin. The solar cells are bonded to the Kapton film. The trim tab panel is of sandwich construction with a lightweight aluminum honeycomb core and thin GFRP faceskins.

Batteries

The two nickel-cadmium batteries provide the power required during launch and ascent phases of the flight (prior to outer solar panel deployment), when in eclipses, and under peak load demands. The battery power is supplied to the primary bus via parallel, redundant battery isolation diodes, redundant battery relays, and the main enable plug. An automatic eclipse-load disconnect reconnect control capability and battery undervoltage disconnect capability are provided both automatic features can be overridden by manual command. The spacecraft load condition in sunlight is such that some equipment must be powered off during eclipse. An automatic load shedding capability is provided which may be enabled and or overridden upon ground command.

Section I Summary

The experiment program was a success in spite of the early anomaly, i.e., the meteoroid shield and solar panel loss which resulted in available power reduction and the non-availability of the solar scientific airlock (SAL) for experiment operations. This success was attributed to the workrarounds that were developed and executed by the ground support personnel and the Skylab flight crews. In addition, the real-time flight planning effort proved to be a significant contribution to the high percent of experimental program accomplishments.

Quick pacts

82-kg (181-lb) cruise stage, 140-kg (309-lb) aeroshell and heat shield Science instruments Mars Volatiles and Climate Surveyor (integrated package with surface imager, robotic arm, meteorology package, and thermal and evolved gas analyzer) Mars Descent Imager Lidar (including Mars microphone) Power Solar panels providing 200 watts on Mars' surface Launch period January 3-27, 1999

The Lockheed Study

The Lockheed team had much to say about the most appropriate type of power source for keeping spacecraft systems operational during a Jupiter mission. Solar panels were ruled out for generating power, due in part to the decreasing insolation (incident solar radiation) as the craft journeyed away from the Sun. Large areas of photovoltaic panels would also present a considerable weight penalty and might be damaged by the high particle fluxes that were anticipated. Because of the distributed mass, it would be more complex to control spacecraft orientation, and the panel array might at times block some of the sensing equipment. Many engineers recommended using radioisotope thermal generators (RTGs) instead, with the hope that the steep costs of such systems would decrease and availability of their isotope fuel would increase by the time construction of the spacecraft commenced. One side benefit of the RTG systems was that the heat they generated could be used to help maintain spacecraft...

Info

SOLAR PANELS (DTM) REMOVED, NO FURTHER TESTS ON PTO REQUIRED. SPECIAL STUB PANELS FROM TCM WILL BE INSTALLED FOR STV TEST. SOLAR ENERGY CONTROLLERS (SECs) REMOVED AND STORED FOR PROTECTION . NO FURTHER REQUIREMENT UNTIL STV TESTS. SEC ACTUATORS SE 4,6,12,AND 14 ALSO REMOVED, RE-INSTALLED ON TEST BRACKETS 5-29-74. SECs AND SEC ACTUATORS RE-INSTALLED ON PTO 6-15-74 FOR STV TEST

Clouds and Climate

Cloud systems regulate variations in incoming solar radiation over Earth's surface and play an important role in distributing solar energy evenly across the surface. As storms move across the planet, excess energy that builds up in the tropics is spread toward the poles. Energy goes back to space from the Earth system in two ways reflection and emission. Part of the solar energy that comes to Earth is reflected back out to space at the same, short wavelengths in which it came to Earth. The fraction of solar energy reflected back to space is called the albedo. Different parts of Earth have different albedos. For example, ocean surfaces and rain forests have low albedos, which means that they reflect only a small portion of the sun's energy -- informally, we think of them as dark. Deserts and clouds, however, have high albedos, and they reflect a large portion of the sun's energy we think of them as light. Over the whole surface of Earth, about 30 percent of incoming solar energy is...

Imager

Senses emitted thermal and reflected solar energy from selected areas of the Earth. It uses a scan mirror system to alternately sweep east to west and west to east perpendicular to a north-to-south path. The rate of scanning allows the instrument to gather data in its five spectral channels while stepping north-to-south from a 1,864 x 1,864-mile area (3,000 x 3,000 km) in three minutes and from an area of 621 x 621 miles (1,000 x 1,000 km) in just 41 seconds.

Thermal Bending

Thermal deflec tions as higli as 0.17 rad were first observed for the 30.5-m 19-03-22A satellite boom which was iinplatec and imperforated. The deflection was significantly reduced in the subsequent l& fTi-'Wn satellite boom by silverplating the outer boom surface such that more solar radiation hitting the side of the boom facing the Sun is reflected, resulting in a more uniform circumferential temperature distribution (ref. 5). Further progress has been made by perforating the boom surface which permits the solar energy to strike directly the side of the boom facing away from the Sun (refs. 8 and IS).

Design Constraints

Extended operation in a space environment before atmospheric entry imposes several material-selection and design problems. The vehicle is exposed to solar radiation, possible meteoroid impacts, a cold-temperature soak, and a vacuum environment. To maintain the desired thermal environment within the vehicle, the thermal protection system is required also to regulate and distribute incoming solar energy. This can be accomplished by means of tailored surface coatings, attitude control, heat sinks, and heat exchangers. The space environment can also alter the chemical and mechanical properties of the heat-shield materials so that their performance in the subsequent entry-heating environment is substantially degraded. Many effects of space operation on material properties can be effectively determined by means of tests in space-simulation vacuum chambers equipped to program variable radiative heating histories.

Science Objectives

Various instruments, notably the infrared spectrometer, will be used to determine the heat balance at Neptune -- the ratio of internal energy emitted to solar energy absorbed. It is known that like Jupiter and Saturn, but unlike Uranus, Neptune emits more heat than it receives from the sun. Measures of the excess energy emitted will have important implications for theories on Neptune's formation and weather mechanisms. Imaging will help characterize wind speeds at different latitudes.

Approach

The ESAS team was tasked to develop new architecture-level requirements and an overall architecture approach to meet those requirements. The architecture requirements developed by the ESAS team are contained in Appendix 2C, ESAS Architecture Requirements. An initial reference architecture was established and configuration control was maintained by the team. Trade studies were then conducted from this initial baseline. In order to determine the crew and cargo transportation requirements, the team examined and traded a number of different lunar surface missions and systems and different approaches to constructing a lunar outpost. A team of nationally recognized lunar science experts was consulted to determine science content and preferred locations for sortie and outpost missions. The use of in-situ resources for propellant and power was examined, and nuclear and solar power sources were traded. The major trade study conducted by the team was an examination of various mission modes for...

Spacecraft

When Genesis' solar arrays are extended in space, the spacecraft resembles an unbuckled wristwatch. The watch's face is the science deck, and the figurative straps are the opened solar panels. The framework of the spacecraft is composed mostly of aluminum, composite materials and some titanium. The use of composites and titanium, lighter and more expensive materials, is an efficient way of conserving mass while retaining strength. Genesis' structure is similar to that used in the construction of highperformance and fighter aircraft.

Criteria

Permanent magnetism in the spacecraft Spacecraft generated current loops Magnetism induced by external fields Currents induced by external fields Additional torques, such as those resulting from the presence of assemblies that rotate relative to each other (e.g., rapidly spinning parts or movable solar panels) should also be considered.

Viking Project

Orbiter power was provided by four solar panels that supplied 620 watts of power at Mars 2 . During the periods when the peak load demand exceeded the power supplied by the solar panels or when the solar panels were not facing the Sun, which occurred during the braking maneuver at Mars, the power difference was provided by two 30 ampere-hour nickel-cadmium storage batteries.

Hand Controller

During operation, heat is stored in several thermal fusible mass tank heat sinks and in the two batteries. Space radiators are located atop the signal processing unit, the drive control electronics, and the batteries. Fused silica second-surface mirrors are bonded to the radiators to lessen solar energy absorbed by the exposed radiators. The radiators are only exposed while the LRV is parked between sorties.

Fig Conclusions

Bay before the SMM servicing could be performed , the SMM deployed solar panels had to be kept clear of the Orbiter radiators and vertical stabiliser during the berthing and re-deployment of the SMM. Had there been the starboard SRMS onboard, the SMM could have been serviced outside the cargo bay in a much simpler manner

Battery Charge

This was not achieved, however, without some abnormal behavior. Factoring out the abnormal experience, the performance corresponded to that predicted. All power system units operated near or above the upper temperature limit of their design characteristics. Figures 81 to 90 present curves of all the power system temperature measurements, as well as solar panel output, voltage, and current, and battery voltage, current, and charge.

Soil Characteristics

The most serious problem in acquiring Mars spectral data is from atmospheric effects on Mars (and on Earth during Earth-based observations) which limit the available wavelength range and attenuate and distort spectral information. Another difficulty is that the composition and grain size of the surface material may not be homogeneous over an area as large as the spatial resolution of the instruments used in remote sensing. Also, the available energy for spectroscopic experiments is limited by the solar constant, the relative position of Earth and Mars, and Martian interaction with solar energy (ref. 122).

Skylab

After approximately 3900 orbits of the earth with 171 days of manned operation from its launch in May 1973, Skylab 1 was a project of unparalleled scientific scope and breadth 1 . There were three visits to Skylab 1 as indicated in Table 5-1 2 . Upon launching Skylab 1, its meteroid shield was torn away from the exterior of the cylindrical workshop along with one of the retracted solar wings. The second solar panel had not been properly deployed, resulting in an overheated and underpowered Skylab 1. The second solar wing was deployed during Skylab 2 mission. The ATM-EPS requirements provided for the ATM to supply electrical power to both the Lunar Excursion Module (LEM) ascent stage and the ATM systems from 24 solar panels via charger battery regulator modules (CBRM). From power conditioning study evaluations, it was shown that a 20 solar module, panel power module configuration was regarded as acceptable. Initially, the solar wing assembly panel layout consisted of 6 solar panels (16...

Earths Energy Budget

For a long time scientists believed that the energy emitted by the sun was constant. The solar constant is defined as the amount of solar energy received per unit surface at a distance of one astronomical unit (the average distance of Earth's orbit) from the sun. Accurate measurements of the variations of the solar constant have been made since 1978. From these we know that the solar constant varies approximately with the 11-year solar cycle observed in other solar phenomena, such as the occurrence of sunspots, dark spots that are sometimes visible on the solar surface. When a sunspot occurs on the sun, since the spot is dark, the radiation (light) emitted by the sun drops instantaneously. Oddly, periods of high solar activity, when a lot of sunspot numbers increase, correspond to periods when the average solar constant is high. This indicates that the background on which the sunspots occur becomes brighter during high solar activity. Equilibrium between the two cavity heat fluxes is...

Permanent Attachment

Single-part dampers are attached either to the main spacecraft structure or to the solar panels. The magnetic hysteresis rods used on the TRANSIT series of spacecraft were inserted into nylon cable clamps which were fastened to the main structure of the spacecraft. The rods that were located in solar array panels, such as in satellites 1965-98A, 1966-76A, and 1967-92A, were inserted into aluminum tubes which were capped off at each end with screws. Both of these methods of attachment permit the rods to be inserted just prior to launch of the spacecraft. This is an important scheduling convenience.

Irtms I Mawds

Solar panel mode When the VO solar panels were illuminated, raw dc power was supplied by the panels. The VO was considered to be in a solar panel mode when the total VO power load demand was within the output power capability of the panels. Share mode A battery solar panel share mode existed when the fully or partially illuminated solar panel output power was insufficient to support the VO power demand. While in this mode, the solar panel output voltage dropped to approximately the battery voltage and entered a high-current operating mode where the solar panel output power was reduced below the nominal output capability. The remainder of the VO power load was supplied by the batteries. An undesirable share mode existed if power demand fell within the solar panel output capability, but the operating point was at the battery voltage. The boost mode was designed to terminate this mode. Battery mode Whenever the solar panels were without illumination the total VO load was supplied by the...

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