1 Formation Flying Rachel Winters Matt Whitten Kyle Tholen Matt Mueller Shelby Sullivan Eric Weber Shunsuke Hirayama Tsutomu Hasegawa Aziatun Burhan Masao Shimada Tomo Sugano

2 Motivation Can enable baseline to form large instruments in space Escort Flights –Provide detection/protection from threats –Provide visual inspection for damage

3 Design A satellite that will fly escort to the space shuttle Satellite provides visual inspection of shuttle exterior for 24 hour period of time Satellite will be transported into space on shuttle Satellite must meet University Nanosat requirements

4 Systems Integration & Management Rachel Winters, Matt Whitten Expendable vs Recoverable spacecraft (90%) Recovery method designed (80%) Determine shuttle-interface requirements (100%)

5 Relative Orbit Control & Navigation Kyle Tholen, Matt Mueller Determine relative orbit to meet mission requirements (90%) Determine major disturbances from orbit and counteract them (100%) Single vs Multiple spacecraft (90%)

6 Configuration & Structural Design Shelby Sullivan, Eric Weber Find general hardware (cameras, thrusters, etc.) (100%) Design structure (material, shape) (90%, pending necessary changes) Solidwork components (60%)

7 Attitude Determination & Control Shunsuke Hirayama, Tsutomu Hasegawa Determine method of attitude control (80%) Single vs Multiple cameras (90%)

8 Power, Thermal & Communications Aziatun Burhan, Masao Shimada, Tomo Sugano Determine power needed by satellite (70%) Battery only vs Solar Cell + Battery (70%) Define thermal environment (outside and inside sources) (80%) Determine insulation needed (60%) Determine transmission method (100%)

9 Trade Studies Expendable vs Recoverable Satellite –method of picture storage –viable method of recovery –reasonable amounts of extra fuel needed Single vs Multiple Satellite(s) –amount of extra fuel needed for plane transfers –ability to “see” entire shuttle with only 1 satellite

10 Solar cells + Battery vs Battery only –Amount of power solar cells can provide in 24 hr period –Amount of power needed by satellite components –Size of battery needed to compliment solar cells vs size of battery needed with no recharge Single vs Multiple camera(s) –Ability to control attitude –Camera size

11 Other Design Aspects Structure: Rectangular satellite with aluminum supports, center of mass designed to be at the center of the prism. Navigation: Will be using DGPS for location and velocity information, magnometer and gyro for attitude determination. Transmission: Decided to store images on memory stick instead of using live transfer.

12 Systems Integration and Management Rachel Winters Matt Whiten

13 SIM Role: Work with all groups to balance workload. Tasks: –Research lightband technology –Perform trade study on attitude sensors –Research ARVD –Research, calculate and design recovery method. Matthew Whitten

14 SIM Attitude Sensors –Distance requires the camera to have the most accurate attitude control –Small satellite requires inexpensive and small equipment Recovery Method –Robotic arm’s length must be able to reach the recovery orbit around the shuttle –Design and format end effect to capture satellite Matthew Whitten

15 Special Requirements Transmission restrictions –NASA operates in the S-band of frequencies, from MHz, the space shuttle is generally contacted at and MHz, and the Orbiter also uses the Ku-band, from MHz. Vibration requirements –Vibration tests with NASA are usually done from Hz.

16 Satellite-Shuttle Interactions Capture feasibility case study –MIR Space capsule –SPARTON satellite –SFU Satellite Automatic movement near to shuttle –Mini AERCam –STS-87

17 Orbital Navigation and Control Group Members: Kyle Tholen Orbit Determination Delta V Estimation –GPS Navigation Matt Mueller Effects of Earth’s Oblatness Propulsion Methods Orbit Modeling in STK

18 Delta V estimation Delta V for orbit transfers estimated with Clohessy Wiltshire equations:

19 GPS Navigation GPS can be used to determine position in orbit Two signals are transmitted from GPS satellites –Precise Position Service (PPS) Very accurate Currently restricted to military applications –Standard Position Service (SPS) Available for anyone to use Not as accurate as PPS

20 GPS Navigation Continued Use Differential GPS (DGPS) for a much more accurate position –Need a known fixed reference position with GPS capabilities –Space Shuttle are GPS certified and position is known very accurately with ground tracking DGPS can potentially be accurate to the centimeter.

21 Orbit Determination Need two orbits to view shuttle from all angles Orbits achieved through small changes in Inclination and Eccentricity

22 Effect Of Earth’s Oblatness Causes secular drift in right ascension, argument of perigee and mean anomaly

23 Earth’s Oblatness Continued Effect on shuttle and satellite nearly the same over 24 hr period deg These values will give the change in the relative distance to the shuttle, estimation of deltaV needed to correct orbit.

24 Propulsion Methods Requirements –Small amount of thrust –Capable of being used numerous times –Small size, light weight –Low price Possible candidates –Small mono-propellant hydrazine thrusters –Cold gas thrusters –Due to simplicity, ease of handling and price, cold gas thrusters were chosen as method of propulsion

25 Orbit Modeling in STK Visualization of relative orbit proved difficult without simulation Created scale simulation of shuttle orbit as well as satellite orbit Useful to visualize relative orbit about shuttle and aid in initial selection of orbit parameters –Use of MATLAB distance function determined final orbit parameters –Simulation proved orbit provided 100% visible coverage of shuttle

26 STK Orbit Simulation

27 Configuration & Structural Design Shelby Sullivan Eric Weber

28 Structure and Configuration Satellite Structure –Cube (60x60x50 cm) –Aluminum Low cost and availability Success on many other satellites Adequate properties for mission Configuration –Keep the moments of inertia near center of cube –Allow space for large camera to see through one face –Allow for proper thermal control

29 Structure and Configuration

30 Structure and Design Gyro Magnometer CPU Transceiver Thruster

31 Structure and Design Future Work –Reconfigure satellite structure to better accomplish design goals –Model remaining hardware –Place selected hardware to accomplish design goals

32 Camera - MegaPlus II EP Megapixel 4872 x 3248 Three sensor grades for “demanding applications” Selectable 8, 10, or 12 bits/pixel “Temperature Resistant” construction

33 Lens - Nikon Super Telephoto 1000mm Angle of view – 2 x 1.4 degrees Length – 24 cm Mass – 2 kg Fixed focal length –Little to no moving parts –Higher vibration resistance –Higher temperature resistance

34 Field of View

35

36 Sample Pictures With Pixels/Meter ~ 360m from shuttl e ~ 700m from shuttle

37 ~360m From Shuttle ~Cross-sectional are of shuttle –400 m^2 Field of View area –105 m^2 ~25% of shuttle captured per photo Accuracy required for view of shuttle –X angle ~ 2.6° –Y angle ~ 0.86°

Meters from Shuttle

39 Attitude Determination & Control Shunsuke Hirayama Tsutomu Hasegawa

40 Why Zero momentum?

41 Moment of inertia of a*b*h cube sat. h b a From Nihon Univ. Text book Once we get angular acceleration, we can get the Moment. Tsutomu and Shunsuke Where, is body frame based moment m ex = Jώ + ω x (Jω)

42 Attitude determination Front ViewSide View xy z

43 Aerodynamic torque Altitude km for worst case S = m 2

44 Gravity-Gradient Torque n 3 = μ = km 3 /s 2 R km 3

45 Solar Radiation Pressure Torque Our surface material is Aluminum 0.02  K  0.04 (surface reflectivity) I s = 1358 w/m 2 at 1 AU

46 Choosing reaction wheel Using Matlab we calculated required torque to change attitude with disturbances. The result is below: Rise Time: Settling Time: Overshoot: % Max Torque: Max torque is 24.6mNm so that we use reaction wheel produced by Sunspace whose max torque is 50mNm. There is error so that we should work on matlab again. For Y axis

47 Problem about simulation Disturbance torque is: Required torques is: We should figure out what is wrong and fix it.

48 Requirement for reaction wheel The rotation speed of satellite should be: 360º/90min = deg/s = rad/min x10 -3 rad/s It takes 90 min to go around the orbit. 360º/90min We use 0.1 rad/s as a rotation speed in matlab

49 Future work calculate a disturbance from magnetic torque. work on matlab with all disturbances.

50 Communications Tomo Sugano

51 Tasks done so far: Communication/CPU selection In-flight Delta V estimation of the mission Atmospheric Drag Analysis Orbital Decay Life

52 FCS and COMM FCS – Flight Control System COMM – Communications (camera is assumed to be part of COMM) Satellite needs to handle both FCS and COMM systems Use of COTS (Consumer Off-the-Shelf) computer(s) aimed COMM utilizes a low-cost COTS transceiver radio

53 CPU selection for the Nanosat Arcom VIPER 400 MHz CPU recommended VIPER is suitable because of its - Light weight, 96 grams - Operable temperature range, -40 C to + 85 C - Windows Embedded feature, easy to program - Computation speed, 400 MHz - Memory capacity, up to 64MB of SDRAM - Embedded audio I/O, necessary for COMM with voice radio Redundancy can be implemented.

54 Arcom VIPER 400 MHz embedded controller

55 Radio selection for the Nanosat Kenwood Free Talk XL 2W transceiver recommended Kenwood Free Talk XL is suitable because of its - COTS nature, low cost - 2W of transmission power, more than enough for non-obstructed space communication, but higher wattage than FRS 500 mW radio - Ability to use both GMRS and FRS frequencies - FRS frequencies recommended because by international treaty FRS (Family walkie talkie) is restricted to 500 mW mW is too weak to penetrate into space - MilSpec cetified

56 Kenwood Free Talk XL 2W FRS/GMRS Transceiver 15 UHF channels (7 FRS and 8 GMRS) 2W output for both categories DC 7.2 V (600mAh) Circuit board weighs only 60 grams Speaker/Microphone/Encapsulation Removed

57 Scheme of FCS/COMM Integration

58 Detailed Scheme of integration

59 Presence of Atmospheric Drag in LEO orbit Atmospheric density is largest at perigee Largest drag is experienced at perigee Atmospheric drag shall be considered if orbit perigee height is <1000 km Atmospheric drag acceleration (D): 1/(AC D /m) is the ballistic coefficient, a measure of resistance to fluid A (projected area normal to flight path) m (mass of spacecraft) f (latitude correction coefficient)

60 Effect of Atmospheric Drag to Orbit Profile Atmospheric drag tends to circularise the probe’s orbit Drag effect greatest at perigee Apogee height consequently reduced Overall altitude is lost unless orbit correction is done Determinant of satellite decay time

61 Drag Coefficient of STS and other LEO probes STS Orbiter (aka the Space Shuttle) STS has a C D of 2.0 at typical mission altitudes in LEO Above 200 km of orbit altitude, use 2.2 < C D < 3.0 Cylindrical probes have larger C D than those of spherical probes Exact C D is hard to predict as LEO environment is not fully understood Currently best determined by actual flight test

62 Consideration of Drag in Formation Flying FF mission is required to last at least 24 hours STS orbiter (primary) typically performs a trim burn once a day Trim burns correct orbit altitude and ascending node Drag differentials present between primary and satellite(s) Possible consideration of LEO drag in our mission

63 Orbital Decay Perturbation in LEO is mainly due to atmospheric drag Orbital decay of space probes (e.g. Space Shuttle, ISS, satellites) Altitude correction “trim burns” necessary to keep probes in orbit Orbit will decay in the absence of trim burns

64 Orbit Lifetime Estimation Estimation of the orbit lifetime of our satellite after mission Consider atmospheric drag effect only Mission orbit is assumed virtually circular for simplicity

65 Orbit Lifetime Equation Circular Orbit Lifetime Equation (Approximation) a 0 = initial altitude S = projected area of the space probe m = space probe mass

66 Exponential Atmospheric Model Scale height, H, obtained from tabulated data

67 Assumptions set forth for our lifetime computation Assumptions: (Made for worst case or shortest decay) m = 50 kg (maximum); S = 0.385m 2 (spherical correction of max volume) C D = 3.0 (upper bound value in LEO probes) a 0 = km (typical altitude for STS or ISS) Δ = 150 – 300 = km (typical re-entry altitude, note the minus sign) f = 1 (ignore latitude effect; not significant (<10%)) ρ 0 = 2.418x kg/m 3 (Table, 300 km base altitude) Unavoidable uncertainty  Scale height, H - Not constant between orbit and re-entry altitude - Take H = 30 km, so β = 1 / (30 km)

68 Computation Result Based on the assumptions we made - T = tau_0 * T = (approx. 1.5 hr of initial orbit period)*(190) = 12 days LEO Nanosat at 300 km of altitude will take 12 days to decay.

69 Conclusion Our Nanosat does not decease for 12 days Retroburn delta-V input to decelerate the Nanosat for faster decay will be costly without a compelling space debris concern(?) Unless allowed to dispose of the Nanosat in space, retrieval is rather recommended(?) Retrieval may be attained fairly easily by using robot arm of STS perhaps equipped with capture net(?)

70 Drag Differential Compensation Different ballistic coefficients between the orbiter and the Nonosat Consequent difference in drag forces exerted during mission Ballistic Coeff. of STS >> Ballistic Coeff. of Nanosat Nanosat must expend Delta-V to keep up with STS orbiter

71 Computations Atmospheric drag acceleration (Da): Drag (acceleration) difference between the two spacecraft: STS: S = 64.1 m 2, C D = 2.0, m = 104,000 kg (orbiter average) sat: S = m 2 (nominal), C D = 3.0 (worst case), m = 50 kg

72 Computations (cont’d) Orbiter speed (assuming circular orbit) Definition of Delta-V (or specific impulse) Mass expenditure of propellant (i.e. GN2 cold gas)

73 Results Using I sp = 65 sec; assume 50 kg for satellite weight Conclusions - At the typical 300 km LEO, Delta-V for 1 day mission is 1.36 m/s - Satellite will need at least 107 grams of GN2 to compensate drag - Besides this Delta-V requirement, we have orbit transfer Delta-V (currently estimated at 1.17 m/s) and ADCS Delta-V.

74 Thermal Control Subsystem Masao Shimada

75 Qs : Direct radiation from the Sun Qe : Radiation from the Earth Qa : Solar radiation reflected back by the earth (Albedo) Qi : Heat generation Qps : Radiation to Space Qpe : Radiation to the Earth Qa Qe Qs Qi Qpe Qps Space Thermal Environment Earth pic:

76 Orbit Model Approximated ISS Circular orbit: Period (T) : Atitude (H) : Shadow time (Ts) : Shadow angle ( ) :

77 o Orbit Model Sunlight Shadow

78 1. Steady-State Approximation Assumptions: 1)Steady State: dT/dt=0 2)Spherical satellite with thermal surface area A= 2.16m^2 so An=0.54m^2 3)Surface characteristic: 4)Heat generration: 50W 5)View Factors: 6) Direct Solar flux: 0 (Cold), 1399w/m^s (Hot) Results: 1)Worst-case HOT: Tmax= K 2)Worst-case COLD: Tmin = K Tmax-Tmin=96.7 K

79 2. Node Analysis QaQsQe QpeQps Thermal Equilibrium Equation Conduction between Node i and Node jRadiation between Node i and Node j

80 Satellite Model for Node Analysis Assume no width for each surface Surface 2 always look downward.

81 Ex: Direct Solar Flux (Worst-case Hot)

82 Worst Cases Worst-case HotWorst-case Cold Earth pic:

83 Surface characteristics Inside of the satellite is painted with L-300 (Black) Conductivity between surfaces : K=0.06

84 Simulink model (Node Analysis)

85 Simulation (Worst-case COLD) Temperatures [K]

86 Simulation (Worst-case HOT) Temperatures [K]

87 Results (Node Analysis) High temperature differences on surface 4 and 5 Use MLI to make thermal disturbances from outside smaller. Need to consider thermal control methods to make temperature higher.

88 Future works Thermal Control by using Thermal Control Elements so that Design Temperature range fits Permissible temperature range of components. More-nodes analysis for accurate simulation

89 EPS Design Trade study of PV-battery vs battery as power source Preliminary analysis (solar array sizing & battery sizing) Power Load Profile Overview of other power susbsystems design - power distribution - power regulation

90 Trade study

91 Trade Study Mission Constraint & Requirement: Length of mission: 1 day Mass <= 30 kg Size: 60 cmx60 cmx50 cm ATITUDE CONTROL Conformal solar array - required spinner to radiate excess heat. Cells not always oriented to the sun, thus reducing power output - for 3 axis stabilized satellite that does not employ active tracking, array’s reduction in output power per total surface area would be approximately 4. Not all surfaces are in the sun. Primary battery - does not affect the choice of attitude control

92 Trade study OPERATING ENVIRONMENT LEO orbit: Worst cold ~-80°C, worst hot ~ 100°C Solar flux variation Radiation PERFORMANCE Conformal solar array - less power output due to cosine loss - single cell efficiencies : 14.8 % (Si), 18.5% (GaAs) - assembled solar array is less efficient than single cell due to inherent degradation, Id ( design efficiencies, temperature, shadowing). Nominal value of Id at life degradation -> ≈1 for short mission (days) - peak power point depends on the array’s operating temperature - required energy storage -> provide power during eclipse

93 Battery - cell voltage decays with Ah discharge - small range of operating temperature -> require thermal control THERMAL CONTROL Both require thermal control, but could be complex for solar array COST Solar array : $ /W. GaAs costs 3 times more than Si $5-$13 per cell Rechargeable battery: $8/cell (NiMh) - $30 (Li Ion) Primary Battery: Lithium type (~ N/A )

94 RISK & SAFETY Solar array - shadowing of one cell results in the loss of entire string. Low risk (with bypass diodes) - minimal safety analysis reporting Primary Battery - limited space qualified battery, safety concerns CONCLUSION Choice of power source depends on power load profile Analysis need to be done to make sure power source meet the mass & area constraint.

95 Primary Battery

96 Solar Cells

97 Rechargeable Battery

98 ANALYSIS Preliminary Solar Array Sizing (according to Space Mission Analysis & Design textbook) Assumption: - Only 2 surfaces will be used to mount solar array. Therefore, optimum available area is 0.72m sq. Maximum number of cells =900 - average power: 50 W - lifetime: 1 day - PPT regulation scheme: Xe=0.6, Xd=0.8 Input: Orbital parameter (ISS orbit) h=300km, inclination = 52 degrees, assume circular orbit => eclipse duration ~36 min, orbital period ~ 91 min Equation: Psa = [( PeTe/Xe) + (PdTd/Xd) P BOL = Po* Id* cosθ P EOL = P BOL * Ld  P BOL because Ld  1 for 1 day mission A sa = Psa / P EOL M a = 0.1 P (with specific 100W/kg) Solar array area, Asa = 0.86m^2 Mass of solar array = 1.18 kg

99 Preliminary array sizing ( according to AEM4332 textbook, pg 495 ) - A array = 1.68 m^2 - N cell = 2100 cells Energy storage sizing ( textbook pg. 485) Mass of battery = 1.55 kg Number of NiCd cells = 22 cells Total mass for solar array + battery = 1.18 kg kg = 2.73 kg Primary battery sizing (lithium sulfur dioxide) Number of cells= 10 Total mass of battery= 6.65 kg (22% )

100 Power Load Profile

101 Power Subsystem General Layout Power source Power distribution Dc-Dc converter Load Energy storage Payload Comm. ADCS Propulsion Thermal

102 Power Distribution & Regulation Main tasks: - power the satellite operation directly - control bus voltage on EPS - control power generated by solar arrays** - charge the secondary battery** Centralize control 28 Vdc bus voltage (regulated) **PPT : extract exact power a satellite require up to array’s peak power Distribution subsystem consist of cabling, fault protection, switching gear, converters (dc-dc) **Battery charging system: Parallel / individual charging

103 Future work Power duty cycle (application profile) - continuous / noncontinuous operation Detail solar array design

104 Thanks, Derek Surka, Joe Mueller