USA ATHENA 1 (LLV 1)

Transcription

1 1. IDENTIFICATION 1.1 Name ATHENA 1 (LLV 1) 1.2 Classification Family : LLV = LMLV(1) Series : LLV = LMLV Version : LLV = LMLV (now ATHENA 1) Category : SPACE LAUNCH VEHICLE Class : Medium Launch Vehicle (MLV) Type : Expendable Launch Vehicle (ELV) 1.3 Manufacturer : Commercial partnership coordinated by LOCKHEED MARTIN Space Systems Division DENVER, COLORADO 1.4 Development manager : LOCKHEED MARTIN Space Systems Division DENVER, COLORADO Telephone: Telex: Vehicle operator : LOCKHEED MARTIN DENVER, COLORADO 1.6 Launch service agency : LOCKHEED MARTIN DENVER, COLORADO 1.7 Launch cost : 16 M$ (1994) 2. STATUS 2.1 Vehicle status : Available 2.2 Development period : First launch : (failure) (success) (1) Lockheed Martin Launch Vehicle December 1997 Page 1

2 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability Low Earth Orbits ORBIT TYPE LEO CIRCULAR SSO CIRCULAR Altitude (km) (Perigee/Apogee) Inclination ( ) Site ESMC WSMC Payload mass (kg) Geosynchronous and Interplanetary Orbits ORBIT TYPE GTO km ; i = 7 INTERPLANETARY Orbit inclination: 4.5 Payload mass (kg) N/A N/A FIGURE 1 - ATHENA 1 PERFORMANCE Data for the 28.5 and 57 incinations are based on the East Range launch at Cape Canaveral. Data for the 90 and 99 inclinations are based on Western Range launch at Vandenberg. December 1997 Page 2

3 The elliptical direct performance shown is for an elliptical orbit with 185 km perigee. One should not try to extrapolate performance to higher altitudes than shown since performance is limited by the amount of hydrazine in the Orbital Adjust Module (OAM). For specific orbits of interest, the ATHENA programme office will run performance calculations. ATHENA orbit injection errors: 3.2 Spacecraft orientation and separation Thermal control manœuvres : Nominal payload separation velocity : Rotation rate : Deployment mechanism type : 3.3 Payload interfaces FIGURE 2 - ATHENA 1 ORBIT INJECTION ERRORS (3 σ) Payload compartments and adaptors Payload fairing description: it's a two-piece clamshell fairing in aluminium with zip-tube separation joints, to avoid contamination. The dynamic envelope represents the diameter and height available to the spacecraft. The weight is 792 kg. The diameter of the fairing is 2.34 m (92 in). FIGURE 3 - ATHENA 1 PAYLOAD ENVELOPE December 1994 Page 3

4 Adaptors - Fixed adaptor section: two sizes of standard adaptor sections. The cm (66 in) diameter is common to all ATHENA applications. Mission peculiar adaptors are designed to this common diameter. FIGURE 4 - FIXED ADAPTOR SECTION - Payload adaptor: a second adaptor section can be added to provide an interface common with the PEGASUS or TAURUS diameter. The payload adaptor ring weight for performance calculations. However, since it is attached to the spacecraft, the spacecraft designer must include the weight in the spacecraft mass properties. * Marmon clamp band separates at two locations. * Springs retract marmon clamp halves. * Separation springs (four places) FIGURE 5 - PAYLOAD ADAPTOR - Secondary payload mounting adaptor is designed to carry secondary payloads. Few satellites require the full performance of either ATHENA 1 or the ATHENA 2. Thus the surplus performance can be used for additional spacecraft carried as secondary payloads. The payloads are located near fairing hinges to avoid fairing motion during separation. December 1994 Page 4

5 3.4 Environments FIGURE 6 - SECONDARY PAYLOAD MOUNTING ADAPTOR Mechanical environment The acceleration experienced by the payload during boost is maximum near the end of the CASTOR 120 TM burn. The acceleration during the ORBUS 21DR burn is lower for heavier payloads (see Figure 7). FIGURE 7 - ATHENA 1 SPACECRAFT ACCELERATION PROFILE (WITH COAST PERIOD) An acoustic resonance exits in the test firings of the CASTOR 120 TM. This is typical for solid motors of this configuration. It occurs for about ten seconds during the first forty seconds of the burn. Based on experience with the Trident first stage motor, the indicated responses are predicted. The specific response for a given spacecraft may be determined with a coupled mode analysis. This data will be refined as the ATHENA structural model is developed. Spacecraft should avoid structural resonances at 55 Hz (see Figure 8). December 1994 Page 5

6 FIGURE 8 - SATELLITE RESPONSE TO CASTOR 120 TM BURN RESONANCE Spacecraft interface design limit load factors envelope the predicted response to quasi-static and transient loads at the centre of gravity of payload with stiffness characteristics. Spacecraft designers should apply factors of safety appropriate to their individual test programmes and risk tolerance to establish design margins (see Figure 9). - Axial load factors envelope payload cg responses to motor ignition transients and steady-state boost accelerations. - Lateral load factors are peak payload cg responses to maximum nozzle deflections during all stages of boost flight. - Axial and lateral load factors should be applied simultanoeusly. - Positive axial load factor indicates tension at the payoad interface. DIRECTION ATHENA 1 Axial Lateral g g g g FIGURE 9 - SPACECRAFT INTERFACE DESIGN LIMIT LOAD FACTORS The acoustic environment induces a random vibration in the ATHENA structure which couples through the payload adaptor interface. These curves, reflecting the two acoustic environments, are the result of conventional modeling with Vibra Acoustic Payload Environment Prediction System (VAPEPS) computer code. December 1994 Page 6

7 FIGURE 10 - ATHENA 1 RANDOM VIBRATION ENVIRONMENT AT SPACECRAFT INTERFACE Acoustic vibrations There are two periods when significant acoustic noise is experienced. The launch environment occurs during the first few seconds of flight resulting from sound reflected off the ground. The transonic environment occurs as the vehicle transit to supersonic velocities. This data was developed from data taken during test firings of the CASTOR 120 TM and flight data from similar sized vehicles scaled to reflect the ATHENA launch configuration. It assumes an empty fairing. The fill factors of individual satellites will increase the amplitude slightly. FIGURE 11 - ATHENA 1 SPACECRAFT ACOUSTIC ENVIRONMENT December 1994 Page 7

8 3.4.3 Shock Measured test data from firings of actual pyrotechnic devices has been scaled to account for the transmission path through the ATHENA structure to the payload interface. The predicted levels resulting from the indicated events are enveloped by the curve. FIGURE 12 - ATHENA 1 SHOCK ENVIRONMENT AT SPACECRAFT INTERFACE Thermal environment Variation of static pressure under fairing Spacecraft compatibility tests In order to preclude dynamic interaction between the spacecraft and the launch vehicle, certain constraints are placed on the spacecraft structural characteristics: to avoid amplification of the spacecraft response to the booster ignition forcing functions, the axial modes of the spacecraft must be greater than 30 Hz, to prevent controllability issues, the lateral modes of the spacecraft must be greater than 15 Hz to provide separation from the first bending mode of the launch vehicle. December 1994 Page 8

9 3.5 Operation constraints ATHENA 1 development schedule FIGURE 13 - ATHENA 1 MASTER SCHEDULE OVERVIEW Ground constraints: Flight constraints: Launch rate capability: Procurement lead time: 4. LAUNCH INFORMATION 4.1 Launch site ATHENA 1 will be launched from the Western Range (Vandenberg AFB) using Space Launch Complex 6 (SLC-6) which was built for Space Shuttle launches. When the idea of Shuttle launchers from the West coast was dropped, the Air Force abandoned the facility. It is foreseen to launch ATHENA 1 from Eastern Range at Cape Canaveral sometime in Stack and Shoot approach The approach launch operations is based on the inherent simplicity and safety of solid rocket motors based launch vehicles. All equipments are checked out before shipping to the launch site. The solid rocket motors are stacked vertically, mated to the interstage and finally to the orbital adjust module (equipment section). The hydrazine for the Orbital Adjust Module (OAM) (equipment section) is added off the launch pad in a ground level fueling facility operation. December 1994 Page 9

10 The spacecraft is processed off the launch site at a payload processing facility. Lockheed provides the payload adaptor which is mated to the spacecraft at some appropriate time. When spacecraft processing is completed the spacecraft is encapsulated with the fairing. The assembly is transported to the launch pad, mated vertically and verified. The timelines are days predicated on an eight (8) hour shift, one shift per day operation. Thus the timelines can be shortened by adding shifts or overtime. An essential point is: Spacecraft are checked-out and made ready for launch Not on the launch vehicle. Miscellaneous last minute tasks, such as removing safe and arm pins, are permitted. Launch support FIGURE 14 - ATHENA LAUNCH SITE OPERATIONS FLOW Simplicity is the key to low cost launch operations. Checkout and launch control equipment is "PC" or personal computer based. The equipment is contained in a van with operator positions for the launch director, the spacecraft director, and the range safety representative. This van and its equipment is used to checkout the Orbital Adjust Module (OAM) at the factory before the equipment is shipped to the launch site. Space is provided for two racks of spacecraft-specific ground equipment. The same van is driven to the launch site. Fiber optic paths and copper wires lead up through the umbilical to the launch vehicle. As the launch rate increases, additional checkout and launch vans are added. FIGURE 15 - LAUNCH SUPPORT December 1994 Page 10

11 4.2 Sequence of flight events Without coast period This illustrates the Hohmann transfer used. The Orbital Adjust Module (OAM) performs a trim burn, as noted, and then the circularization burn at apogee. The various events of the ATHENA 1: 780 km circular orbit, 108 indirect injection are listed hereunder. FLIGHT SEQUENCE FLIGHT TIME (s) EVENTS Ignition of first stage Maximum dynamic pressure (1.44 b) Maximum static acceleration (6.9 g) Second stage ignition Nose fairing separation Maximum static acceleration (6.1 g) OAM V trim burn Burnout Orbit insertion burn Burnout Nota: the flight time, given here as an example, depends on the mission. December 1994 Page 11

12 Without ascent coast period FIGURE 17 - ATHENA 1 INDIRECT INJECTION PROFILE The various events of the ATHENA 1: 463 km circular orbit, 28.5 inclination are listed hereafter. FLIGHT TIME (s) EVENTS Ignition of first stage Burnout of first stage Ascent coast Fairing separation Ignition of second stage Burnout of second stage Transfert coast Ignition OAM Burnout OAM 4.3 Launch record data LAUNCH DATE NUMBER OF SATELLITES ORBIT RESULT REMARK LEO Failure Vandenberg LEO Success Vandenberg LEO Success Cape Canaveral LEO Success Kodiak Island December 2001 Page 12

13 Failures LAUNCH DATE RESULT CAUSE The ATHENA flew east when it was supposed to fly south. Self destruction command issue D by USAF range safety officer 160 s into the flight Two anomalies led to failure : the 1 st in 1 st stage began at 5 s. The burning hydraulic fluid recirculated back up under skirt and became too hot for instruments, causing at 60 s to pitch up and down. Previsional reliability : CASTOR 120 (99.9%); ORBUS 21 DR (99.9%) Success ratio : 75% (3/4) next occured at 127 s when 2 s electrical short caused by internal pressure build up caused to lose the IUM reference. 4.4 Planned launches 2002: no data available 5. DESCRIPTION 5.1 Launch vehicle FIGURE 18 - ATHENA 1 OVERVIEW 5.2 Overall vehicle Overall length Maximum diameter Lift-off mass (approx.) : 18.3 m : 2.34 m (92 in) : t December 2001 Page 13

14 5.3 General characteristics of the stages STAGE 1 2 OAM Designation CASTOR 120 TM ORBUS 21 DR OAM Manufacturer LOCKHEED UTC-CSD OLIN (1) Length (m) Diameter (m) Dry mass (t) Propellant: Solid Solid Solid Type Class 1.3 Class 1.3 Storable Fuel HTPB 86% HTPB Hydrazine Oxidizer TP-H1246 UTP B Propellant mass (t): Fuel Oxidizer TOTAL Tank pressure (bar) Total lift-off mass (t) (1) Formerly known as Rocket Research Upper part DESIGNATION VEHICLE EQUIPMENT BAY FAIRING Manufacturer OLIN LOCKHEED Mass (t) Launch vehicle growth: yes, ATHENA 2 ATHENA 3 December 1997 Page 14

15 5.4 Propulsion STAGE 1 2 OAM Designation CASTOR 120 TM ORBUS 21 DR OAM Engine CASTOR 120 TM ORBUS 21 DR ASC(1) Manufacturer THIOCKOL UTC-CSD OLIN Number of engines axial 6 control Engine mass (kg) Feed syst. type Mixture ratio Chamber pressure (bar) Cooling Albative Albative - Specific impulse (s) Sea level Vacuum Thrust (kn) Sea level Vacuum x x 6 Burning time (s) Nozzle expansion ratio Restart capability No No Yes (1) Altitude Control System 5.5 Guidance and control Guidance The inertial navigation unit is the Litton LN 100 L (laser gyro). December 1994 Page 15

16 5.5.2 Control The guidance processor is built by Lockheed STAGE 1 2 Flight coast and orbit adjust Pitch, yaw Gimballed EMA - TVC OAM Roll OAM OAM OAM Deflection ± 5 ± EMA = Electro-Mechanical Actuator A cold gas attitude control system is under consideration as a lower cost option for missions not requiring precise orbits 6. DATA SOURCE REFERENCES 1 - Lockheed Launch Vehicles, October 26, Aviation Week, May 2, 1994, p Space News, January 31 - February 6, AIAA Low-cost liquid upper stage for the Lockheed Launch Vehicle 5 - AIAA Lockheed Launch Vehicles - An effective use of reliable, low cost propulsion 6 - Flight International September 1994, p Jane's Space Directory IAF-94 S Lockheed Launch Vehicle solid propulsion systems 9 - AIAA A stable of space motors ready for the expanding ELV market December 1994 Page 16