Power subsystem Microsatellite

Power sub transcription Micro orbiterThis subsystem is responsible for supplying motive to the entire air, converting solar cell energy to on-board battery energy, and distributing office to various other subsystems.The power subsystem of the micro transmit is designed for a remote sensing military com armorial bearing to carry out on solarise-synchronous orbits at 700 km altitude at an inclination of 98.19 degrees. The payload of the microsatellite includes a multispectral remote sensing camera which takes picture of polar region in a visible spectrum and a surrey GPS receiver especi everyy designed for low ball orbit. Microsatellite payload weighs 5 kg and with a mean power consumption of 9W. Sub-system power budget is estimated according to the payload power requirement with 15 percent margin. Total estimated power requirement for the microsatellite is 70W.Microsatellite subsystem Power Allocation - End Of Life Estimated Microsatellite Power 70 WSubsystem % of Operating Po wer Power (W)Payload 1510.5LSTS BusPropulsion 00Thermal Control107Attitude Control1510.5Power1510.5Communications2014C D Handling 107Structure 00Margin1510.5Total10070The power subsystem of the microsatellite is designed for Low Earth Orbit for five years period. The power estimated for subsystem has a 15% contingency margin. Primary power source for the satellite is the solar cast that is body mounted on the microsatellite. The satellite is in near polar sun-synchronous orbit at an altitude of 700 km, total orbital period of the satellite is 98.77 min. The microsatellite experience eclipse for about 35.29 min. solar array for the microsatellite is designed according to the delegating requirement. Batteries are secondary power source during the eclipse when no sun light is available. The selection of the solar cell and batteries are made according to power required end of life of the satellite and trade study surrounded by different solar cell and batteries but decision is made to satisfy the estimated mass size and power budget of the satellite. As the satellite is a cube shaped and spins stabilized body mounted solar panels to places on all the four sides of the satellite for a uninterrupted supply of power to the subsystems.Altitude 700.00kmEarths r6,378.00kmTotal Power Requirement (const. day and night)70.00WattsEarths Gravitational Constant 3,98,600.00km2/s2Power transfer efficiencies-Xd0.85Xe0.75Inherent Degradation Id0.80Worst Case( deg)23.00(deg) cathexis Life (yrs)5.00(Yrs)Life time Degradation (Ld)0.98Angle (rad)1.12(rad)Angle (deg)64.30(deg)orbital Peroid (P) (sec)5,926.21(sec)Maximum Eclipse Peroid (Tn) (sec)2,117.08(sec)Minimum Power Sunlight (Td) (sec)3,809.12(Sec)Average Solar array power (Psa) (W)134.23WMultijunction Solar (GainP/GaAs) Po 301.00W/m2BOL Power (Pbol)221.66W/m2EOL Power Requirement (Peol)216.17W/m2Solar array Area (m2)0.62m2 push-down storage of Solar Array (kg)3.36kgSolar array lean ( body mounted so Msa x 4 )13.42kgThe primordial power source of the microsatellite is chosen to be Multijunction Solar cells (GainP/GaAs). These solar cells submit an efficiency of 23 percent and most advanced for their category. The required solar panel area of the microsatellite to sufficiently clog the power requirement of the microsatellite subsystem is 0.62 m2 but for a body mounted microsatellite, all the four faces of the cube shaped satellite will have the following area. The estimated weight of the solar panels is 3.4 kg and the total weight of all the panels on the satellite is 13.5 kg. The main returns of the body mounted solar panels is such that they have much life expectancy as they are not exposed to radiation for a long time, but it is compensated with the additional weight of the solar panels. The primary power source should be able to gene dictate 135 Watts of power to sustain the power requirement of the subsystems as well as enough to press the batteries as they are the secondary power source of the kick.For Given Ni H cell Assuming Data for 700 km altitude Energy parsimony 100.00W.h/kgDOD1.60 0.27 log10 cyclesPower during Eclipse70.00WAltitude700.00kmBattery Voltage28.00VoltsXb-l0.90No .of eclipes per day15.005 year MissionOrbital Peroid (P) (sec)5,926.21SecTime of Night (Tn) (sec)2,117.08SecEb (energy supplied during eclipse) (W.h)45.74W.hCycles26,607.25Depth Of Discharge (DOD)0.411a)Ebcap (energy battery capacity required) (W.h)112.87W.h1b) Battery Capacity (A.h)(assuming voltage is 28 v)4.03A.h2. Total Battery Mass (kg)1.13KgThe secondary power source is required to generate power during eclipse in the orbit to sustain microsatellite subsystems. The secondary power source for the mission is chosen to be NiH batteries as they are good for long cycle life and they have advantage of mass and volume over most of the ongoing batteries available. They have good specific energy density of 50 W.hr/kg. The main advantage of these batteries is such that they are widely use d in space mission and constantly updated with new technologies. They have depth of discharge of 40% that is good for this kind of mission. Total secondary power source weight is 2.3 kg.(((((((((( References SMAD and System consolidation Aegis))))))))Communication subsystem The communications subsystem is the lead for the interface between the satellites and the ground stations. The communications subsystem helps in demodulating the received uplink subscribes and transmitting downlink signals .The subsystem also helps us to maintain a bounce back over the satellite by transmitting received range tones and by playacting as logic between receive and transmitted signals.Data orderThe remote sensing microsatellite is designed for a Low Earth orbit at an altitude of 700 km. The payload of the satellite is a multispectral camera that takes picture of the poles in visible spectrum. The 20 degrees minimum elevation angle and a resolution of 50 is simulated for the satellite and the da ta rate is calculated for the satellite.Altitude (km)700.00Radius of Earth (km)6378.14Orbit Peroid (mins)98.77Ground Velocity ( km/s)6.76Node Shift (L = S) (deg) 24.76 (deg)20.00 (deg)57.86Zc27818.52Za133.06Z3701467.63DR (Visible)(bps)37014676.33Maximum Time in View (min)6.66The data rate calculated is 37Mbps adding 10 percent margin data required to send back to ground station is estimated to 40Mbps.Band Link TechnologyFor the current microsatellite mission an S-Band telecommunication system is researched, analyzed, and chosen as the best system for establishing communication between satellite and the ground station.ApplicationSpecificationsDownlink Rate Max 2.5MbpsPower RF Output .4WPower Consumption3.4WWeight420gVolume190X135X22 mm3The table above shows the specification if the Surrey Satellite S band communication system transmitter details. This has an advantage of low mass, power and data rate which completely satisfy the mission constrains. The above transmitter system also h as a S-Band overture for this transmitter which has specifications as follows.(((((((((((((((((((memo com2 // surrey satellite))))SpecificationsNumber of Antennas Needed43dB Beamwidth 35Weight80gVolume82X82X20 mm3Link Budget Link budget for the system S band communication system is designed considering the factor to transmitting the data rate of 40Mps within 6.5mins or 400 sec.The link budget is a process of accounting all the workable gains and losings during transmitting and receiving the signals from transmitter to receiver.The equations below are used to check over link budgetTotal spacecraft received power (uplink budget)Uplink Signal to sound ratio (Will help determine probability of bit error)Total Ground Station received Power (downlink budget)Downlink Signal to Noise ratio (Will help determine probability of bit error)2.4.1 Slant RangeThe Slant range was calculated as follows for a 5 degree elevation angle.2.4.2 Attenuation of the SignalThe biggest contributor to the a ttenuation of the signal is free space loss. There are many other losses such as cable loss, polarization loss, cloud, rain, etc.The absolute frequency used for the S-Band calculation is 2.2GHz.Atmospheric loss is caused by absorption due to such factors as oxygen and water vapor in the atmosphere. Atmospheric, rain, clouds and ionosphere scintillation were assumed to be 0.5dB for 2.2GHz. Further investigation into these effects needs to be completed next semester. With X-Band the total loss due to these factors was calculated to be 0.76dB. S-Band is expected to have a much lower loss.Polarization loss was estimated from 92.4.3 Calculating EIRPThere will be passive losses in the equipment such as losses in the coax cables. This number was used from the previous year.Power transmitted was obtained from the specification on the Surrey transmitter as 0.4 Watts.Looking at the Co-Polar gain on Figure 2 it is seen that there is a gain of at least 0dB for angles between +/- 70.2.4.4 Groun d Station Antenna GainUsing an antenna that is 4.5m in diameter with efficiency of 0.55 the gain is calculated as follows2.4.5 Signal to Noise CalculationThe signal to racquet ratio will determine the objet dart Error Rate (BER), as determined from the following graph 8.From this graph it can be seen that to obtain a Bit Error Rate of 10-5 which is ordinary of space missions, a signal to noise ratio of 4.4 dB is needed.The Link Budget calculations will determine if the system will meet the 4.4 dB of signal to noise ratio at the ground station.System Noise is a function of temperature and was determined from table 13-25 2.4.8dB is above the minimum 4.4dB theoretical signal to noise ratio required. This leaves only a 0.4dB margin which needs to be approved upon. The output RF power could easily be increased from 0.4Watts by use an amplifier, but would be at the expense of the satellite power budget. The Surrey Satellite equipment is a viable solution.Thermal Subsystem The thermal control subsystem is the integral part of the satellite design. It helps out all the components that are exposed to thermal environment are not affected badly. Thermal control subsystem accomplish safe working of all the satellite subsystems and their components by constituting a thermal model.The following process includes inputs from different subsystem of the satellite by identifying the thermal loads that will acting on them during the mission lifetime as well as their operating tempertature for the smooth running of the mission.Thermal LoadsThe satellite experience or exposed to thermal enviroment during gound testing, transportation, pitch , orbit transfer and operational orbits. The thermal environment concerned is during its operation in space. There are four main loads acts on the satellite during its mission.(smad)Direct Solar Radiation The main source of direct solar radiation is the Sun. It is major source of environmental heating on the satellite, it is a perpetual en ergy source and it constant to the fraction of second. The intensity of the sunlight on the earths mean distance of 1 Astronomical unit (AU) is 1367 W/m2. Earths albedo Albedo is the reflected sunlight reflected from earth . It is highly as it is the fraction of incident sunlight that is refected back to space. Refletivity increases over land rather than in oceans. Reflectitivy increases with decreasing local solar -elevation angle.Earths Infrared Energy It is also refereed as blackbody radiation, all incident sunlight do not reflected back as abledo rather earth absorbs it and re-emit it as IR (infrared Energy ) or blackbody radiation.Free Molecular heating This load is result of the bombardment of the individual molecules present in outer reaches of the atmosphere. It affects during the lance ascent of the satellite. The thermal control susbsystem is designed for a sun synchronous Low Earth Orbit at an altitutde of 700km and at an inclination of 98.19 degrees.The main verbal exp ression in designing the thermal control system is to first define the worst case hot (maximum loads) and worst case cold (minimum loads ) acting on the satellite in the orbit and the opertonal temperature operational and survival temperature of each component installted