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I-Cone
Launch Vehicle |
Stack using I-Cone |
Teaming with Saab-Ericsson, a study was completed to evaluate the innovative I-Cone® launch concept. There is an increasing demand from payload and small satellite providers to obtain space access in a timely and affordable way. The trend is that launch vehicles tend to get bigger while the existing small launch vehicles have a high weight-to-orbit cost. Today, specialized solutions exist for payloads to gain access to space; these solutions require a high degree of coordination. The European ASAP and the U.S. ESPA are methods to avoid the specialized solutions for auxiliary payloads. I-Cone is a step further than the current open secondary launch methods.
I-Cone is a system to provide fast, frequent, flexible, and affordable access to space for small payloads, either micro / nano satellites (dispenser function) or payloads (satellite platform function). The I-Cone replaces the satellite adapter during launch carrying the primary manifested space vehicle (SV) loads and uses the normally empty adapter volumes for payload and equipment.
The I-Cone baseline provides the standard EELV interface to the launch vehicle and a standard separation interface to the primary space vehicle. New launch vehicle or space vehicle interfaces are not required to include I-Cone on an EELV launch. Impacts to the primary space vehicle are minimal and unnoticeable. The modular approach of the I-Cone design also allows for implementation on other launch vehicles including Sea Launch and Delta II.
 Communication I-Cone |
 0.5m Diameter Optical Telescope |
 Active Dispenser |
The other scenario involving the I-Cone concept is the dispenser⁄mothership case. This configuration allows microsats or nanosats to be deployed from the I-Cone adapter as required. Accommodation of these miniature vehicles is provided both on the interior and exterior of I-Cone. Within the dispenser concept, there are two variations. The first variation is an I-Cone dispenser with no internal avionics, only the capability to deploy the micro⁄nanosats. Variation two consists of an I-Cone mothership that may transport micro ⁄nanosats to a different parking orbit for later deployment. |
Another option currently being evaluated is a combination of micro⁄nanosats in addition to small payloads on a single I-Cone.
Overall, I-Cone provides a reduced-cost, modular option for access to space. A secondary benefit of I-Cone is the fact that the I-Cone payloads, whether microsats or instruments, are transparent during the integration process. The launch integration flow is not hindered by the inclusion of an I-Cone as opposed to a mechanical adapter.
I-Cone is a concept poised to revolutionize U.S. methods for access to space. Not only will I-Cone provide lower launch costs, but also I-Cone payloads may be operational almost immediately following separation from the launch vehicle.
| Item |
Proposed Baseline Implementation |
| Mission Life |
Baseline designed for minimum of 1 year with consumables for 2+ years |
| Launch Vehicle |
Standard 1666 separation interface in baseline approach with optional 1194, 1663, or 937 interfaces available; 1575 bolted interface to launch vehicle included |
| Profile |
Worst-case orbit parameters used in evaluation to bound system design; On-board propulsion system provides for launch insertion errors, small inclination changes, drag makeup, and deboost, Δ V=410 m⁄s with 20% margin |
| Launch Readiness |
Proposed 24 months for initial unit readiness for payload and 18 months for follow-on units |
| Payload Mass |
Accommodated by the I-Cone inner structure and potential exterior mounted shelves |
| Space Vehicle Mass |
Maximum launch mass limited only by primary mission requirements and launch vehicle, 506 kg baseline estimate |
| Payload Power |
28 VDC, four (4) switched services with over and under current protection |
| Space Vehicle Power and Margin |
EOL 300 W total system load prediction including 15% margin; (1) - 21 Ah battery max 33% DOD during worst eclipse; 2.6 m2 deployable solar array is driven at orbital rate by single axis actuator |
| Thermal |
Passive, cold-biased thermal design consisting of dedicated spacecraft and payload radiators; robust autonomous thermostatic control of resistive heaters |
| C&DH Architecture |
Time tagged and event driven commands. Telemetry and I⁄O for payloads along with discrete analog⁄digital I⁄O |
| Payload Data Rate |
Payload data rate accommodated: 48 kps with 1553 I⁄F and 2 Mbs with RS-422 I⁄F plus 16 kbs for SV housekeeping data |
| Data Storage |
Standard DHU configuration can accommodate 5 Gbit and is expandable to 16 Gbit |
| CMD Interface |
Both 1553B and RS-422 serial I⁄F provided |
| GN&C (3 sigma values) |
Three Axis Stabilization employed: Zero Momentum (3 RWA), Gyro and Fixed Head Star Tracker, Pointing Accuracy of
10-50 arcsec achievable with SIRU upgrade |
| Communication |
S-Band System, Omni Antenna; 2 Kbps Uplink; 2 Mbps Downlink Data Rate provides > 6db Link Margin; six 8 minute contacts⁄day provides 25% margin relative to budgeted 60 minute downlink time |
| Radiation |
All avionics are resistant to >30 Krads dose and are SEU tolerant |
| Magnetic Cleanliness |
2.5 A-m2 Peak, Single-Axis Transient SV Control Magnetic Moment and a S⁄C Harness E⁄M Dipole of 0.72 A-m2
at the
S⁄C Body Outer Surface are Worst Case Estimated Environments from SV |
| Redundancy and Fault Tolerance |
SV bus is single string; some functional overlap and selective redundancy allows for increased fault tolerance. The over-voltage protection circuitry of the PCE provides for protection against exceeding the maximum bus voltage specifications and a degraded mode of operation for battery charging |
Contact
space@atk.com
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