TAURUS. 2.2 Development period : ; (commercial version)

Transcription

1 1. IDENTIFICATION 1.1 Name 1.2 Classification Family : Series : Version : 2110/2210* Category : SPACE LAUNCH VEHICLE Class : Small Launch Vehicle (SLV) Type : Expendable Launch Vehicle (ELV) 1.3 Manufacturer : Orbital Sciences Corp. (OSC) Atlantic Boulevard DULLES, VIRGINIA Tel. : (703) Fax : (703) Development manager : Orbital Sciences corp. (OSC) 1.5 Vehicle operator : U.S. Air Force Space System Division and OSC (commercial missions) 1.6 Launch service agency : Orbital Sciences Corp. (OSC) 1.7 Launch cost : M$ (FAA estimation) 2. STATUS 2.1 Vehicle status : Operational 2.2 Development period : ; (commercial version) 2.3 First launch : (SSLV version) ; (commercial version) * A SSLV (Standard Small Launch Vehicle) with the Peacekeeper ICBM first stage was used for the maiden launch. On all subsequent flights the first stage is a Castor 120. A four-digit nomenclature is used to identify configurations (example 2210): 1st - Type of first stage (1 = Peacekeeper also referred as ARPA ; 2 = CASTOR 120), 2nd - Fairing size (1 = 1.6 m fairing; 2 = 2.3 m fairing), 3rd - Third stage motor (1 = ORION 38; 2 = STAR 37), 4th - Stage 4 (0 = none). December 2000 Page 1

2 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability Low Earth Orbits ORBIT TYPE LEO CIRCULAR PEO CIRCULAR SUN SYNCHRONOUS CIRCULAR Altitude (km) (Perigee/Apogee) Inclination ( ) Site Cap Canaveral (Eastern range) Vandenberg (Western range) Vandenberg Payload mass (kg) Taurus (1) - 960(2) 740 Taurus (1) - 770(2) 580(1) - 600(2) (1) Launch from North Vandenberg (2) Launch from South Vandenberg Geosynchronous orbits 400 to 450 kg GTO capacity can be achieved with the standard configuration. There are several options for increasing the available mass-to-orbit (using stretched ORION motors for Stages 1 and 2 or replacing ORION 38 Stage 3 motor with a spinning STAR 37FM motor). Injection accuracy Orbital injection errors are dominated by variations in the Stage 3 total impulse, navigation errors, and other stage performance parameters. Three-sigma injection accuracies are summarized in the following table. ERROR TYPE Injection Apse (km) Non-injection Apse(*) (km) Mean altitude (km) Inclinaison ( ) THREE-SIGMA TOLERANCE ± 10 ± 50 ± 30 ± 0.15 (*) OSC refers to the "non-apsidal injection" technique since the upper stage motor fires at a point that is neither apogee nor perigee December 2000 Page 2

3 FIGURE 1 - PERFORMANCE TO 28.5 LEO ORBITS (CCAFS) NVAFB: launches from North Vandenberg SVAFB: launches from South Vandenberg FIGURE 2 - PERFORMANCE TO 90 LEO ORBITS (VAFB) December 2000 Page 3

4 FIGURE 3 - PERFORMANCE TO SUN SYNCHRONOUS ORBITS (VAFB) 3.2 Spacecraft orientation and separation Thermal control manœuvres Nominal payload separation velocity Rotation rate Deployment mechanism type : yes : mission specific (typically about 0.6 m/s) : < 15 rpm : spring release 3.3 Payload interfaces Payload compartments and adaptors Payload fairing description: offers two payload fairings: 1.6 m (63 in.) and 2.3 m (92 in.) configurations. OPTION 1.6 m CONFIGURATION 2.3 m CONFIGURATION Structure Composite sandwich Composite sandwich Material Graphite-epoxy facesheets over aluminium honeycomb Graphite-epoxy facesheets over aluminium honeycomb External length (m) Usable length (m) External diameter (m) Mass (kg) Sections 2 2 December 2000 Page 4

5 The two halves of fairing are structurally joined along their longitudinal interface using a low contamination frangible joint system. At separation, a gas pressurization system is activated to pressurise the fairing deployment thrusters. The fairing halves then rotate around external hinges that control the fairing deployment to ensure that payload and launch vehicle clearances are maintained. Payload adaptor and separation: The separation system is a marmon clamp band design that employs two aluminium interface rings that are clamped by dual, semi-circular stainless steel clamp bands with aluminium clamp shoes. Separation velocity is provided by up to 8 matched spring actuators that impart up to 37.6 Joules of energy. Payload access provision: One access door is provided for the payload as a standard service. For the 1.6 m fairing, the size is 305 mm x 305 mm; for the 2.3 m fairing: 457 mm x 610 mm. The access doors are RF-opaque. Additional doors can be provided as non-standard services. 3.4 Environments FIGURE 4 - PAYLOAD FAIRING OPTIONS (1.6 m AND 2.3 m DIAMETER) Mechanical environment Design limit load factors due to the combined effects of steady state, low frequency transient and quasisinudoidal accelerations are defined in the table below. These values include uncertainty margins. Note that loads produced by pressure oscillations generated during solid rocket motor combustion, known as resonant burn, produce a response that is both transient and quasi-sinusoidal in nature. December 2000 Page 5

6 Lift-off EVENT Subsonic Resonant Burn Transonic Supersonic Gust S1 Resonant Burn AXIAL, g steadystate ± oscillating (resultant) ± 3.5 (- 6.0/+ 1.0) ± 4.0 (- 7.0/+ 1.0) ± 4.5 (- 8.0/+ 1.0) ± 0.5 (- 4.5/- 3.5) ± 2.0 (- 6.0/- 2.0) LATERAL (g) ± 2.0 ± 2.0 ± 2.5 ± 2.5 ± 1.5 Motor Burn Out ± 0.5 Vibrations Sine vibration Figure 5 defines the maximum payload interface sinusoidal vibration levels associated with the lift-off (10-25 Hz), Stage 0 resonant burn (45-65 Hz) and Stage 1 resonant burn (65-75 Hz) forcing functions. Also indicated is the lower bound on the range of responses predicted for payloads to date. Random vibration Figure 6 defines the maximum payload interface random vibration levels as well as the lower bound on responses predicted for payloads to date. Also indicated is an envelope of the flight measured data. December 2000 Page 6

7 FIGURE 5 FIGURE 6 RANGE OF SINE VIBRATION LEVELS PAYLOAD INTERFACE VIBRATION LEVELS Acoustic vibrations Fairing interior noise levels for the 1.6 m and 2.3 m diameter fairings are defined in figure 7 and figure 8 respectively. Levels are shown with a maximum fill factor included and with no fill factor included. flight microphone data shows that peak acoustic environments occur at lift-off and near the point of maximum dynamic pressure. As noted in the figures, acoustic levels are dependent on spacecraft geometry (fill factor) and, to a lesser extent, upon acoustic absorption characteristics. As a standard service, missionspecific analyses will be conducted by Orbital using a customer provided statistical energy model of the payload. December 2000 Page 7

8 3.4.3 Shock FIGURE 7 FIGURE 8 MAXIMUM FLIGHT LEVEL PAYLOAD MAXIMUM FLIGHT LEVEL PAYLOAD ACOUSTIC ENVIRONMENT ACOUSTIC ENVIRONMENT (1.6 m diameter fairing) (2.3 m diameter fairing) The maximum predicted shock levels at the payload interface are presented in figure 9. These levels are applicable to a payload using ' standard payload separation system. FIGURE 9 - PYROSHOCK AT THE PAYLOAD INTERFACE (SRS: SHOCK RESPONSE SPECTRUM) December 2000 Page 8

9 3.4.4 Thermal environment Ground operations: upon encapsulation within the fairing and for the remainder of ground operations, the payload environment is maintained by the Environmental Control System (ESC). Fairing inlet conditions are selected by the customer, and are bounded as follows: - dry bulb temperature : C controllable to ± 2 C of setpoint, - dew point temperature : 3-17 C, - relative humidity : determined by dry bulb and dew point temperature selections and generally controlled to within ± 3%. Powered flight: the maximum fairing inside wall temperature will be maintained at less than 121 C, with an emissivity of This temperature limit envelopes the maximum temperature of any component inside the payload fairing with a view factor to the payload with the exception of the Stage 3 motor. The maximum upper stage motor surface temperature exposed to the payload will not exceed 177 C, assuming no shielding between the aft end of the payload and the forward dome of the motor assembly. Whether this temperature is attained prior to payload separation is dependent upon mission timeline. Fairing deployment will be initiated such that the maximum free molecular heating rate is less than W/m² Cleanliness level The payload processing area is in a class clean room environment. An ECS from payload encapsulation through vehicle lift-off is provided with a filter unit to give class air conditions. 3.5 Operation constraints Ground constraints: co-ordination is exercised by OSC. Launch rate capability: 4 per year (planned); up to 12 per year (capability). Procurement lead time: 24 months. December 2000 Page 9

10 USA Ø Integration process The typical mission cycle for a specific launch is detailed in Figure 10. FIGURE 10 - TYPICAL MISSION INTEGRATION SCHEDULE Ø Launch operations: The upper stages are integrated horizontally in the MAB (Main Assembly Building) prior to their shipment to the launch site. Within 3 weeks prior to launch, the integrated hardware is transported from the MAB to the launch site in 3 major pieces: the Stage 0 motor, the Stage 0/1 interstage and the upper stage assembly. Upon arrival at the site, the Stage 0 motor and Stage 0/1 interstage are immediately erected onto the launch pad and the upper stage assembly is positioned inside the integration tent in preparation for final vehicle testing. December 2000 Page 10

11 Once connections and flight simulation are successfully completed, is "ready" for payload transport. This typically occurs at L-12 days. At the Site, the ECE (Encapsulated Cargo Element) is cranelifted from transporter, rotated to a horizontal position, and placed next to upper stage assembly inside the integration tent. Following the successful completion of tests, the ECE is mechanically mated to the upper stage assembly. Using a special handling fixture and 2 cranes, the upper vehicle stack with ECE is hoisted, rotated to vertical and mated to previously erected Stage 0. Following final vehicle testing, is armed and the pad is cleared for launch. These activities, which typically commence on L-1 day and conclude by L- 12 hours, bring the vehicle to a launch-ready configuration. This timeframe also represents the final period for payload access under normal situations. 4. LAUNCH INFORMATION 4.1 Launch sites Both VAFB and CCAFS facilities are available to users. Wallops Flight Facility can also be used. Vandenberg Air Force Base (VAFB), California - Location: 34.7 N/120.6 W Launches are conducted from the Space Launch Complex SLC 576E located on the coast of North VAFB. Cap Canaveral Air Force Station (CCAFS), Florida - Location: 28.5 N/81.0 W Launch Complex 46 (LC-46), proven commercial facility will be used for inclinations between 28.5 and 40. Wallops Flight Facility (WFF), Virginia - Location: 38.0 N/75.3 W Due to its inherant range safety launch azimuth limits, this Site could be used to provide inclination only between 38 and 55. Launch vehicle processing Figure 11 depicts the typical flow of hardware from the factory to the launch site. December 2000 Page 11

12 FIGURE 11 - HARDWARE FLOW - FACTORY TO LAUNCH SITE December 2000 Page 12

13 Transportable Launch Support Equipment The system also includes a complete set of transportable Launch Support Equipment (LSE) designed to allow to be operated as a self-contained satellite delivery system. This LSE includes the launch stand, Launch Equipment Van (LEV), Launch Support Van (LSV) and assorted mechanical and electrical ground support equipment. The LSV serves as the actual control center for conducting a launch and includes consoles for Orbital, range safety, and payload personnel. While the system is capable of self-contained operation, it is typically launched from an established range; thus, the vehicle and LSE are designed for compatibility with existing US Government ranges at Vandenberg Air Force Base (VAFB), Cap Canaveral Air Station (CCAS) and Wallops Flight Facility (WFF). A communications network connects the LSV with the Range Control Center (RCC). 4.2 Sequence of flight events FLIGHT TIME (s) EVENTS Stage 0 ignition and lift-off Stage 0 burnout, separation and Stage 1 ignition Stage 1 burnout and separation Stage 2 ignition Fairing separation Stage 2 burnout; Stage 2/3 coast Stage 2 separation and Stage 3 ignition Stage 3 burnout/orbit insertion Payload separation December 2001 Page 13

14 4.3 Launch record data FIGURE 12 - TYPICAL DIRECT ASCENT MISSION PROFILE (circular orbit inclined at 97.6 ) LAUNCH DATE NUMBER OF SATELLITES ORBIT RESULT REMARK LEO Success WTR - MX Stage LEO Success WTR - Castor 120 Stage LEO Success WTR LEO Success WTR LEO Success WTR LEO Failure WTR Failures : one LAUNCH DATE RESULT CAUSE At T + 83 s, as the stage 2 lit, the vehicle went out of control about 5 s before righting itself and continuing on. Because of the performance loss, orbit was not achieved and satellites were lost The problem was focused on Thrust Vector Control (TVC): the actuator device on the stage 2 steering system did not move December 2001 Page 14

15 Previsional reliability : - Success ratio : 83,3% (5/6) 4.4 Planned launches For 2002, no data available; three launches have been scheduled by DESCRIPTION 5.1 Launch vehicle FIGURE 13 - VEHICLE CONFIGURATION December 2001 Page 15

16 5.2 Overall vehicle Overall length Maximum diameter Lift-off mass (approx.) : 27.9 m : 2.38 m : 73 t 5.3 General characteristics of the stages STAGE Designation Manufacturer Thiokol Alliant Alliant Alliant Length (m) Diameter (m) Dry mass (t) Propellant: Type Solid Solid Solid Solid Fuel HTPB-AI* HTPB-AI* HTPB-AI* HTPB-AI* Oxidizer NH 4 CIO 4 NH 4 CIO 4 NH 4 CIO 4 NH 4 CIO 4 Propellant mass (t) Fuel Oxidizer TOTAL Tank pressure (bar) Total lift-off mass (t) * HTPB 88% (classe 1.3) Upper part DESIGNATION VEHICLE EQUIPMENT BAY FAIRING SPELTRA SYLDA 5 Manufacturer Litton OSC/Vermont Composites (1.6 m) OSC/R-Cubed Composites (2.3 m) Mass (t) 360/400 kg December 2000 Page 16

17 Launch vehicle growth As an optional service, the ORION 38 Stage 3 can be replaced by a spinning upper stage using Thiokol's STAR 37FM motor. Uprating options are the XL (same stretched stages) and XLS (with two strap-on motors). Ballistic phase capability: In a typical direct ascent mission, the coast time can reach up to seven minutes. With a typical parking orbit mission profile, the ballistic phase duration is above sixteen minutes. 5.4 Propulsion STAGE Designation CASTOR 120 CASTOR 50 S-G ORION 50 ORION 38 Engine Manufacturer Thiokol Alliant Alliant Alliant Number of engines Engine mass (kg) Feed syst. type Mixture ratio Chamber pressure (bar) Cooling Ablative Ablative Ablative Ablative Specific impulse (s) Sea level Vacuum Thrust (kn) Sea level Vacuum Burning time (s) Nozzle expansion ratio Restart capability No No No No December 2000 Page 17

18 5.5 Guidance and control Guidance Litton LR-81 Inertial Measurement Unit (strapdown) aided by position data from Trimble GPS receiver. Avionics system is an all-digital distributed processor design Control FLIGHT PHASE STAGE Pitch, yaw (Deflection) Gimballed nozzle (hydraulic) Gimballed nozzle Gimballed nozzle Gimballed nozzle Roll None N 2 cold gas RCS N 2 cold gas RCS N 2 cold gas RCS 6. DATA SOURCE REFERENCES 1 - Castor 120 motor: development and qualification testing results - J. HILDEN - B. POIRIER - AIAA Paper n September JANE'S Space Directory International Reference Guide to Space Launch Vehicles - S.J. ISAKOWITZ - J.P. HOPKINS Jr, J.B. HOPKINS - AIAA Edition 4 - Launch System Payload User's Guide - OSC - September 1999, Release December 2000 Page 18