Dione Solar Array SOW 7-29-20.pdf

DIONE SOLAR ARRAY STATEMENT OF WORK

1. Objective/Requirements

To purchase a gimbaled, deployable solar array with release mechanism(s) that meets the specifications in this document. This hardware will be used for Dione, which is a 6U CubeSat.

2. Characteristics, Scope, and Specs

The contractor shall provide the following:

2.1. One (1) array with two (2) deployable wings and applicable release mechanism(s)

2.1.1. See Figure 1 for the notional design. Note that the wings are symmetric about the gimbal axis of rotation.

Figure 1 – Notional Design

2.2. Consumables for hold down release mechanisms (HDRMs) for five (5) deployments. If applicable, the contractor shall assume that a deployment requires energizing all primary deployment components (e.g. resistor or nichrome wire) or all secondary deployment components (not both). The contractor shall assume that all resistors must be replaced after being energized once. The lifetime of any nichrome wire shall be negotiated after award. See Section 3.14 for more details about HDRMs and deployment.

2.3. Vacuum-compatible mating connector set, including one spare

2.4. Protection diodes

2.4.1. Protection diodes may be incorporated into the wings or on a separate board to be installed inside the spacecraft bus

2.4.2. See Section 3.18 and 3.19 for detailed requirements

3. The system shall meet the following requirements:

3.1. Array shall comply with the following mechanical constraints

3.1.1. Two 6U sides are available for stowed array wings

3.1.2. Each 6U side can accommodate 364mm x 209mm x 10mm maximum stowed volume

3.1.3. Gimbal mechanism shall fit within 100mm x 100mm x 10mm volume on 2U side

3.1.4. Each HDRM utilizes a thermal knife, where a resistor or nichrome wire heats enough to melt a high-strength polymer

3.1.5. All solar cells shall point in the same direction when deployed, within five (5) degrees

3.1.6. HDRMs and any support structure for stowing the wings shall be confined to the 6U sides of the spacecraft.

3.1.7. Support structure shall not block the DESA instrument field of view (zenith and nadir) or the star tracker field of view (zenith). See Figures 2, 3, and 4. A solid model is available upon request. Note that fields of view extend to infinity, even though they have a finite size in the model.

3.1.8. The wings shall be of a size and shape such that, when deployed, they can be rotated to a position where they are fully out of the DESA, GRIDS, NMS and Star Tracker field of view (FOV). Note that fields of view extend to infinity, even though they have a finite size in the model. Note that staying out of the GRIDS Small FOV is required. Staying out of the GRIDS Large FOV is preferred.

3.1.9. Mechanical design shall adhere to CubeSat standards for spacecraft using a rails-style deployer

Figure 2 – Spacecraft Isometric View

Figure 3 – Spacecraft Anti-RAM Face (Zenith to the Left)

Figure 4 – Spacecraft Nadir Face (RAM Direction to the Right)

3.2. Array shall include a one-axis gimbal with the axis of rotation as shown in Figure 1.

3.3. Gimbal range of motion shall be at least +/- 175°

3.4. Gimbal shall be designed for a minimum lifetime of 10,000 cycles. Rotation from -175 degrees to +175 degrees and back to -175 degrees is one cycle.

3.5. Maximum gimbal slew rate shall be at least 12 mrad/sec.

3.6. Gimbal power draw at maximum slew rate shall be no more than 5W

3.7. Inrush current to the gimbal mechanism at startup shall be no more than 4A

3.8. Array mass shall be no more than 1.5kg. This is for the entire array, including but not limited to the wings, gimbal, wiring, support structure and HDRMs.

3.9. Array shall produce at least 84W at Beginning of Life (BOL) and at least 69W at End of Life (EOL)

3.10. Array shall be designed for a minimum of 1-year mission lifetime

3.11. The two deployable wings shall be designed to meet the following temperature ranges:

3.11.1. Survival temperature range: -100°C to +110°C

3.11.2. Operational temperature range: -90°C to +100°C

3.12. The gimbal and associated electronics, also known as the solar array drive assembly (SADA) shall be designed to meet the following interface temperature ranges:

3.12.1. Survival temperature range: -50°C to +90°C

3.12.2. Operational temperature range: -40°C to +70°C

3.13. Array shall be designed for circular, LEO orbit (for radiation degradation purposes)

3.13.1. Altitude = 400 – 600km

3.13.2. Inclination = 70° – 85°

3.14. Array shall be compatible with Ibeos (formerly Cubic Aerospace) 150W EPS

3.14.1. Produces a total solar array voltage per string within 19V-40V during the mission lifetime

3.14.2. All strings will be connected in parallel to a single battery charge regulator

3.14.3. EPS is Peak Power Tracker (PPT)

3.14.4. Cells produce less than 6 Amps

3.15. The array shall meet the following requirements with respect to the HDRMs

3.15.1. Each wing is stowed with no more than two HDRMs

3.15.2. Redundant deployment components for each HDRM, such as two resistors

(primary and secondary) for each HDRM, are preferred.

3.15.3. If the HDRM(s) do not have redundant deployment components, the vendor shall provide safe actuation parameters, such as temperature, current and time, for ground testing in air and vacuum, to ensure that the deployment circuity will not be stressed or damaged during ground testing.

3.15.4. Array wiring allows all primary deployment components to be heated via a single input circuit

3.15.5. If the design includes secondary deployment components, array wiring allows all secondary deployment components to be heated via a single input circuit

3.15.6. Circuit to heat primary deployment component(s) requires no more than 2.5A

3.15.7. If the design includes secondary deployment components, circuit to heat secondary deployment component(s) requires no more than 2.5A

3.15.8. No need to send array back to manufacturer to reset HDRM(s) after deployment test

3.15.9. HDRM at technology readiness level (TRL) 8 or better

3.16. The array shall be back-wired, to minimize the magnetic dipole when the cells are illuminated.

3.17. Environmental acceptance test shall include a minimum four (4) thermal cycles, thermal cold/hot deployment (whichever is worst case, based on torque margin) and vibration

3.18. The array shall include a protection diode for the spacecraft batteries. This diode may be incorporated into the wings or on a printed circuit board separate from the array.

3.19. If a separate printed circuit board is used for the battery protection diode, the board shall meet the following requirements:

3.19.1. Include a blocking diode for the Ground Support Equipment (GSE) used to charge the batteries

3.19.2. Include one connector each for the solar array, Ibeos EPS and GSE

3.19.3. Use smart diodes, which minimize thermal dissipation

3.19.4. Conform to footprint and mounting location requirements to be negotiated after contract award

3.20. Each solar cell shall be accompanied by a bypass diode

3.21. The array shall have a microcontroller interface, used to control the gimbal.

3.22. The microcontroller interface shall provide angular position feedback for the gimbal accurate to +/- 1 degrees or better.

3.23. Each wing of the array shall have an integrated temperature sensor. This means the array will have a total of two temperature sensors.

3.24. EEE parts for the gimbal shall be capable of surviving a radiation total dose environment of 7 krad behind 1mm of aluminum. The vendor shall assume no shielding is provided by the spacecraft bus structure. If the gimbal housing provides less shielding than the equivalent of 1mm of aluminum, the total dose requirement shall be increased as appropriate.

3.25. Solar cell interconnect tabs, EEE parts and printed circuits on the wings, gimbal electronics and diode board (if applicable) shall be coated with low outgassing space flight materials. Exceptions to this requirement shall be limited to connector pins, connector sockets and connector mating surfaces. Excepted features shall be free of conformal coating.

3.26. Vendor shall provide calculations of voltage and power at BOL and EOL.

Calculations shall be performed under worst-case conditions. For example, BOL with no losses at cold temperature may yield worst-case high voltage. EOL with losses at hot temperature may yield worst-case low voltage and power.

3.27. Losses accounted for in power calculations shall include, but not be limited to, the following:

3.27.1. Coverglass and adhesive degradation (darkening due to total non-ionizing dose)

3.27.2. Solar cell degradation (due to total ionizing fluence)

3.27.3. Assembly

3.27.4. Harness

3.27.5. Testing errors

3.27.6. Blocking diode

3.27.7. Bypass diode

3.27.8. Contamination

3.27.9. Thermal cycling

3.28. Vendor shall provide results from solar cell power tests

3.29. Vendor shall supply instructions for customer to perform a solar cell power test on the finished array

3.30. Vendor shall supply instructions for customer to stow array and reset HDRMs

3.31. Vendor shall supply a schedule with proposal

3.32. Vendor shall supply, with proposal, examples of past performance with respect delivery schedule

3.33. Note that vendor proposals will be scored as follows:

3.33.1. 50% of score: meeting technical performance requirements

3.33.2. 20% of score: cost

3.33.3. 20% of score: delivery schedule, including supporting examples per 3.32

3.33.4. 10% of score: personnel qualification

4. Place of Performance Vendor facility

5. Period of Performance Delivery of flight array and all other deliverables a maximum of 10 months after award of contract