PSLV 1. IDENTIFICATION. 1.1 Name. 1.2 Classification Family : SLV Series : PSLV (1) Version : PSLV-C (2)

Transcription

1 1. IDENTIFICATION 1.1 Name 1.2 Classification Family : SLV Series : (1) Version : -C (2) Category : SPACE LAUNCH VEHICLE Class : Medium Launch Vehicle (MLV) Type : Expendable Launch Vehicle (ELV) 1.3 Manufacturer : ISRO (Indian Space Research Organization) VSSC (Vikram Sarabhai Space Centre) TRIVANDRUM KERALA LPSC (Liquid Propulsion Centre) THIRUVANANTHAPURAM Telephone : (91) / Fax : (91) Development manager : ISRO/Department of Space (Indian Space Research Organization) Antariksh Bhavan New BEL Road BANGALORE Telephone : (91) Fax : (91) Vehicle operator : ISRO SHAR Centre Sriharikota Range ANDHRA PRADESH Telephone : (91) Fax : (91) Launch service agency : INDIA Antrix Corporation Ltd New BEL Road BANGALORE Telephone : (91) /274 Fax : (91) antrix@isro.ernet.in 1.7 Launch cost : about 20 M$ (1) Polar Satellite Launch Vehicle (2) -Commercial; a previous version was -D (Development) December 2000 Page 1

2 2. STATUS 2.1 Vehicle status : Operational 2.2 Development period : First launch : (Failure) (-D) (Success) (-D) (Success) (-C) 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability The performance capability has been steadily improved from 820 kg in its first development flight (-D1) in 1993 to kg into 817 km polar orbit in 1997 (-C1). The last data available are given in the table Low Earth Orbits ORBIT TYPE LEO CIRCULAR SSPO CIRCULAR Altitude (km) (Perigee/Apogee) Inclination ( ) Site SHAR (1) SHAR SHAR Payload mass (kg) (1) Sriharikota Range Geosynchronous Orbits About 850 kg, at 18 inclination, GTO capability can be achieved with the current configuration and kg should be reached with the next -C3 version. The launch vehicle has no capability for a direct GEO injection Injection accuracy Three-sigma injection accuracies are summarized for a 817 km SSPO in the following table. ORBITAL PARAMETERS Apogee/Perigee Inclination Velocity correction ( V) ERROR ± 35 km ± m/s December 2000 Page 2

3 INDIA FIGURE 1 - -C1 PAYLOAD CAPABILITY FOR PLANAR AND POLAR LAUNCHES FIGURE 2 - PAYLOAD VS INCLINATION FOR DIFFERENT CIRCULAR ORBITS 3.2 Ø Ø Ø Ø Spacecraft orientation and separation Thermal control manœuvres Nominal payload separation velocity Rotation rate Deployment mechanism type December 2000 : : 0.8 m/s : : spring release Page 3

4 3.3 Payload interfaces Payload compartments and adaptors The fairing is an aluminium alloy structure fabricated in two halves. It consists of a spherical nose cap and a 20 conical section at the forward end, a long cylindrical section and a short conical frustrum boat tail. The conical sections are stiffened half shell structures and the cylindrical section is an integrally stiffened isogrid structure made up of 3 panels of 1.5 m height each: length : 8.3 m, primary diameter : 3.2 m, mass : kg. FIGURE 3 - PAYLOAD ENVELOPE FOR For LEO missions, multiple satellites will have to be accommodated to exploit the full capability and utilise the available payload envelope within the fairings. Development of a smaller fourth stage with a propellant loading of 1 t (L1) and a single engine is planned to achieve optimum payload volume. Figure 4 presents the various payload configuration options on. December 2000 Page 4

5 Payload fairing description FIGURE 4 - PAYLOAD ENVELOPE CONFIGURATION OPTIONS The payload adaptor is a framed structure with a conical shape. The spacecraft separation system is incorporated on the forward end ring. Payload adaptor assembly has a height of 510 mm. Its diameters at aft and forward ends are respectively mm and mm. The structure is able to support a kg spacecraft. Suitable adaptors can be designed to meet users requirement for multiple spacecraft launch. Separation is provided by a combination of a Merman clamp, a pair of explosive bolt cutters and zip cord. Four springs mounted on the payload adaptor provide separation velocity. Payload access provisions Removable doors can be provided in the heatshield to permit limited access to the spacecraft following fairing installation. Standard cut-out access size are 540 x 540 mm. Change in size and location of the cutouts are coordinated with ISRO during finalization of the agreement. December 2000 Page 5

6 3.4 Environments Mechanical environment The maximum static and dynamic accelerations occurring at spacecraft interface during each stage are given in the following table. MAXIMUM VEHICLE ACCELERATION LEVELS STAGE LATERAL (g) LONGITUDINAL (g) STEADY DYNAMIC STEADY DYNAMIC I II ( *) 0.2 ( *) III IV Response due to N 2 O 4 depletion * Response due to UDMH depletion Vibrations - Sine vibration The sinusoidal qualification and acceptance test levels given in the table are applied at the base of the spacecraft. These flight levels are anticipated levels and are multiplied by a factor 1.5 for spacecraft qualification. This factor is intended to ensure a reasonable degree of margin. QUALIFICATION AND ACCEPTANCE TEST LEVELS OF SINUSOIDAL VIBRATION Longitudinal axis FREQUENCY RANGE (Hz) Lateral axis QUALIFICATION TEST LEVEL (zero to peak) 6.75 mm (DA) 1.80 g 1.8 g g 3.75 g 3.75 g g 0.75 g 6.75 mm (DA) 0.67 g 0.45 g DA: Double Amplitude Nota: criteria for notching will be decided after coupled analysis. ACCEPTANCE TEST LEVEL (zero to peak) 4.5 mm (DA) 1.2 g 1.2 g g 2.5 g 2.5 g g 0.5 g 4.5 mm (DA) 0.45 g 0.30 g December 2000 Page 6

7 Random environment QUALIFICATION AND ACCEPTANCE TEST LEVELS OF SINUSOIDAL VIBRATION FREQUENCY (Hz) Overall level Duration Acoustic vibrations QUALIFICATION PSD g 2 /Hz g (rms) 2 min/axis ACCEPTANCE PSD g 2 /Hz g (rms) 1 min/axis The most severe acoustic environments occur at lift-off due to jet noise and in transonic flight due to unsteady shocks and boundary layer noise. The high frequency dynamic excitations in the payload area are generated mainly by acoustics. The estimated acoustic levels are given in the table below Shock QUALIFICATION AND ACCEPTANCE TEST LEVELS OF SOUND PRESSURE OCTAVE BAND SOUND PRESSURE LEVEL IN db CENTRE FREQUENCY (Hz) QUALIFICATION ACCEPTANCE Overall level in db Duration min min Shock levels depend on spacecraft mass and payload adaptor construction. A typical shock spectrum is shown in Figure 5. December 2000 Page 7

8 3.4.4 Thermal environment FIGURE 5 - TYPICAL PAYLOAD SEPARATION SHOCK SPECTRUM Prelaunch environment inside heatshield AIR FLOW RATE kg/h Air circulation velocity 2 m/s max Temperature Adjustable between 15 C and 25 C Relative humidity 40 to 60% Filtration Class Variation of static pressure inside heatshield Adequate venting has been provided in the heatshield to ensure that the difference in static pressure inside and outside the heatshield is less than 10 kpa. The variation of static pressure with time is given in Figure 6. December 2000 Page 8

9 3.5 Operation constraints Ground constraints FIGURE 6 - PRESSURE VARIATION INSIDE HEATSHIELD Due to the geographic features of the coast line and the location of SHAR range, the launch azimuth for a polar launch has to be limited to a maximum of 140. Further, the flight safety procedures like limiting the Instantaneous Impact Points (IIP) of the vehicle away from the coast and overflying the land mass are followed as per the norms. Thus, for the nominal flight trajectory, the IIP trace is at least 372 km away from any land mass. Also the flight sequence is so designed as to have the impact points of the separated stages outside the economic zones of the countries and also away from the dense shipping lanes. In case of unacceptable deviations from the nominal flight corridor, the vehicle will be destroyed by telecommand. In order to achieve the final orbital inclination of 99, the vehicle, launched South-East from SHAR, has to perform a yaw manoeuvre of about 55. The yaw manoeuvre is initiated at T s based on the flight safety constraints, among other considerations. Launch rate capability 1-2 per year Procurement lead time About 26 months Integration process The vehicle integration and launch operations are currently the direct responsibility of ISRO. Participations by industries in these areas are expected to progressively increase. While the metallic motor structures are fabricated by industries, the propellant casting and processing them into segments ready for stacking is accomplished at Solid Propellant Booster Plant (SPROB) located in Sriharikota Launch Range (SHAR). The liquid propellant tankages and light alloy interskirts supplied by Indian aerospace industries are equipped and the integrated liquid stages are delivered by Liquid Propellant Systems Centre (LPSC). The vehicle avionics systems and the equipment bay are realised at Vikram Sarabhai Space Centre (VSSC), which is the lead centre responsible for launch vehicle design, development and integration tasks. December 2000 Page 9

10 The composite motor cases, nozzle assemblies, ignitors, pyro systems and other auxiliary system elements are also realised at VSSC. All the interskirt subassemblies and Vehicle Equipment Bay are fully integrated and checked out at System Integration Facilities (SIF), VSSC Valiamala Complex (VMC) which is located 30 km off Thiruvananthapuram. These modules are transported by road in specific container-trailor systems to the launch complex at SHAR, where the final vehicle integration takes place. Figure 7 represents schematically the various workcentres and the vehicle hardware flow sequence. Launch operations FIGURE 7 - WORKCENTRES AND PRODUCTION FLOW is integrated on the launch pad. The sequence starts with the placement of vehicle baseshroud assembled to nozzle end segment on the launch table. The PS-1 booster segments are stacked making the interface clevis joints on pad, followed by the attachment of six strap-on solid motors. The second stage is moved to the launch pad in transtilt trailer then elevated to vertical orientation, lifted and placed above the booster stage. The third and fourth stages are brought to the pad as a prestacked module and hoisted into position to complete the vehicle assembly. Integrated vehicle checks are carried out before mating the spacecraft. The closure of fairings is followed by the final checkout and launch countdown. Current occupancy of vehicle on pad is about 50 days and typical launch count down begins at T - 45 h. December 2000 Page 10

11 4. LAUNCH INFORMATION 4.1 Launch site launches are performed from the ISRO facilities at Sriharikota (refer to ASLV data sheet). The Sriharikota High Altitude Range (SHAR) is located in South-East India along the Bay of Bengal (13.73 N/80.24 E), at about 80 km North of Madras. The location of SHAR centre imposes severe launch window and range safety constraints on launches into polar orbits. Launch azimuth is limited to 140, requiring polar-orbit missions to be launched in a southeasterly direction, followed by a 55 yaw manoeuvre. Similarly, low inclination orbits are restricted to inclinations of about 40 or higher, unless yaw manoeuvres are used. A launch corridor at an azimuth of 102 is available for GTO missions, with a resulting inclination of 18. The SHAR launch centre includes one launch complex for the, which consists of a pad, a fixed umbilical tower and a 75 m tall mobile service tower for assembly of the rocket. The mobile service tower has a clean room for payload integration onto the launch vehicle. Motor casting and vehicle pre-integration testing are performed at other facilities at the SHAR centre. The launch complex has also been modified with cryogenic fuelling and servicing capabilities to support upcoming GSLV launches. FIGURE 8 - LAUNCH FACILITIES LOCATION December 2000 Page 11

12 FIGURE 9 - LAUNCH COMPLEX Payload processing Effective integration of the spacecraft to requires timely preparation. The operations at launch pad are carried out in the 3 phases. The spacecraft campaign duration at SHAR is normally 30 days before launch. The sequence of operations on the launch pad is described in Figure 10. Working days T Spacecraft preparation Checkout validation Phase-1 (SP-1) Spacecraft preparation Phase-2 (SP-2) Filling & final S/C operations Spacecraft activities at MST Phase-3 (SP-3) S/C Assembly & C/O Heatshield assembly Phase-3 checks with S/C Launch preparation Count down FIGURE 10 - TYPICAL SPACECRAFT OPERATIONS SCHEDULE December 2000 Page 12

13 Launch vehicle processing The upper stage module stacking, satellite mating to vehicle and close of fairing are followed by about 45 h of prelaunch countdown when the propellant servicing, arming and checkout operations are carried out before lift-off. The vehicle checkout operation is remote through fibre optic link from checkout computer located in LCC, 5 km away from the pad. From T - 10 min, the Automatic Launch Sequence (ALS) takes over the final phase of checkout, culminating in the ignition command to the booster stage at T Sequence of flight events The main events in the typical flight sequence for a sun-synchronous polar mission are as follows (refer to Figure 11). FLIGHT SEQUENCE TIME AFTER (s) LIFT-OFF EVENTS 4 solids + stage 1 ignition 2 solids ignite at 2.5 km 4 solids burnout and separation at 24 km 2 airlit solids burnout and separation at 41 km Stage 2 ullage motors ignition, stage 1 separation + stage 1 retro motors ignition Stage 2 ignition at 76.7 km Fairing separation at 123 km Closed loop guidance begins Stage 2 shutdown, separation and retro motors ignition Stage 3 ignition at 279 km Stage 3 separation at 579 km Stage 4 ignition at 701 km Stage 4 shutdown at 826 km Satellite injection Nota: the flight time, given here as an example, depends on the mission. 4.3 Launch record data LAUNCH DATE NUMBER OF SATELLITES (*) Partial failure of the fourth stage ORBIT RESULT REMARK LEO Failure PEO Success PEO Success PEO Success (*) -C1 PEO Success -C2 PEO Success -C3 GTO Success -C4 LEO Success -C5 December 2003 Page 13

14 Failures FIGURE 11 - TYPICAL FLIGHT SEQUENCE FOR SUN-SYNCHRONOUS POLAR LAUNCH LAUNCH DATE RESULT CAUSE The upper rocket section and the satellite fell back in the sea An error in the software which failed to correct a flightpath anomaly after 3 rd stage separation Previsional reliability : - Success ratio : 87.5% (7/8) 4.4 Planned launches Three more launches are planned, one flight/year. December 2003 Page 14

15 5. DESCRIPTION 5.1 Launch vehicle FIGURE 12 - VIEW OF 5.2 Overall vehicle Overall length : 44.4 m Maximum diameter : 5.1 m with strap-on motors (fairing: 3.2 m) Lift-off mass (approx.) : 294 t December 2000 Page 15

16 5.3 General characteristics of the stages STAGE Designation PSOM (1) PS-1/S125 PS-2/L37-5 PS-3/S7 PS-4/L2 Manufacturer ISRO ISRO ISRO ISRO ISRO Length (m) Diameter (m) Dry mass (t) Propellant: Type Solid Solid Liquid Solid Liquid Fuel HTPB-AI HTPB-AI UDMH HTPB-AI MMH Oxidizer NH 4 CI O 4 NH 4 CI O 4 N 2 O 4 NH 4 CI O 4 MON-3 (2) Propellant mass (t) Fuel Oxidizer TOTAL 9 x Tank pressure (bar) Total lift-off mass (t) - - Helium - Helium 11 x (1) Strap-On Motors (2) MON-3: Mixed Oxide of Nitrogen (3% NO and 97% N 2 O 4 ) Upper part DESIGNATION VEHICLE EQUIPMENT BAY FAIRING Manufacturer ISRO Inertial Systems Unit Hindustan Aeronautics Ldt Mass (t) Launch vehicle growth GSLV is an upgraded version of : the six solid strap-on boosters are replaced with four liquid strapon similar to the stage 2. A cryogenic upper stage replaces the last two stages of. December 2000 Page 16

17 5.4 Propulsion STAGE Designation PSOM PS-1/S125 PS-2/L37.5 PS-3/S7 PS-4/L2 Engine - - VIKAS (1) - - Manufacturer VSSC (2) VSSC LPSC (3) VSSC LPSC Number of engines 1 x (4) Engine mass (t) Feed syst. type - - Turbopump - Pressure Mixture ratio Chamber pressure (bar) Cooling Ablative Ablative Regenerative Ablative Regenerative Specific impulse (s) Sea level Vacuum Thrust (kn) Sea level Vacuum 662 x x 2 Burning time (s) Nozzle expansion ratio Restart capability No No No No Yes (1) Viking-class type (2) Vikram Sarabhai Space Centre (Solid propulsion group) (3) Liquid Propulsion Systems Centre (4) One upgraded engine for launch vehicle growth 5.5 Guidance and control Guidance Guidance: IGS (Inertial Guidance System), located in Vehicle Equipment Bay surrounding stage 4 base. A Redundant Strapdown Inertial Navigation System (RESINS) is designed with 3 dry-tuned gyros in a skewed configuration and 4 servo-accelerometers, which feed the navigation processor providing data every 500 ms to the guidance and control processor that issues the steering commands. December 2000 Page 17

18 5.5.2 Control STAGE Pitch, yaw (Deflection) - 1 nozzle with SITVC (1) Nozzle gimbal Flex-seal nozzle gimbal Electromechanical nozzle gimbal Roll Additional control by SITVC on 2 strap-on RCS 2 thrusters supplied by hot gas RCS of 4 th stage Nozzle gimbal + RCS after MECO Deflection - - ± 4 ± 2 ± 3 (1) SITVC: Secondary Injection Thrust Vector Control 6. DATA SOURCE REFERENCES 1 - User's Manual - ISRO - Issue 3, May presentation - Published by ISRO Head Quarters - September JANE'S Space Directory International Reference Guide to Space Launch Vehicles S.J. ISAKOWITZ - J.P. HOPKINS Jr - J.B. HOPKINS - AIAA Edition 5 - Management of launch vehicle mission in a multiple technology environment as applied to Indian satellite launch vehicles Dr. K.S. RAO - IAF Paper 98 - A Flight experience of Indian polar satellite launch vehicle S. RAMAKRISHNAN - S. SRINIVASAN - IAF Paper 98 - V Vibroacoustic environment simulation for qualification of an integrated subassembly of S.A. PALANISWAMI - S.V. SHARMA - IAF Paper 99 - I The Indian Polar Satellite Launch Vehicle () - A low cost launcher for small satellite missions E. JANARDHANA - V. MANOHARAN - IAF Paper 99 - V Passenger payload on -C2: the first commercial launch service by Antrix/ISRO S. RAMAKRISHNAN - K. RAMACHANDRAN - IAF Paper 99 - V Indian upgrades and variants for multi-mission role - S. RAMAKRISHNAN - S.S. BALAKRISHNAN - IAF Paper 00-V107 December 2000 Page 18