SOYUZ M$ (US Department of Transportation estimate) Commercial mission costs to be quoted

Transcription

1 1. IDENTIFICATION 1.1 Name 1.2 Classification Family : A Series : A-2/SL-4(*) Version : A-2/SL-4(**) Category : SPACE LAUNCH VEHICLE Class : Medium Launch Vehicle (MLV) Type : Expendable Launch Vehicle (ELV) 1.3 Manufacturer : Ts SKB 18, Pskovskaya Str. SAMARA Russian Federation Telephone : (78462) Fax : (78462) Development manager : Ts SKB 1.5 Vehicle operator : Ts SKB 1.6 Launch service agency : RKA (Russian Space Agency) 3, Miusskaya Square MOSCOW Launch cost : M$ (US Department of Transportation estimate) Commercial mission costs to be quoted (*) A-2 : Charles Sheldon of US Library of Congress designation. The other operational version in the series is the A-2-e/SL-6 called MOLNIYA SL-4 : DOD designation (**) An upgraded version designated U2 was introduced in 1982 and used synthetic kerosene (SinTin) in the core stage to provide higher performance. The U2 was flown 70 times, all successfully from 1982 through This vehicle was then discontinued awing to the end of production of its SinTin fuel. The initial designation of SEMYORKA for A-class vehicles has been used for the ICBM versions of the first launch vehicles December 2002 Page 1

2 2. STATUS 2.1 Vehicle status : Operational 2.2 Development period : "A" series launch vehicles were derived from SS-6/SAPWOOD ballistic missiles. was developed from VOSKHOD for the purpose of launching the manned spacecraft. U version was introduced in First launch : November 16, 1963 (VOSKHOD) November 28, 1966 (SOYOUZ) 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability Low Earth Orbits ORBIT TYPE LEO CIRCULAR LEO ELLIPTICAL SUN SYNCHRONOUS Altitude (km) (Perigee/Apogee) / / Inclination ( ) Site BA BA PL PL BA PL PL Payload mass (kg) BA: Baikonur PL: Plesetsk Geosynchronous and Interplanetary Orbits No capability Injection accuracy For a standard orbit (altitude 200 km; inclination 51.8 ): Period of revolution Inclination : ± 22 s : ± 3.5 min December 2002 Page 2

3 FIGURE 1 - PERFORMANCE CAPABILITIES IN CIRCULAR ORBITS FROM BAIKONUR FIGURE 2 - PERFORMANCE CAPABILITIES IN CIRCULAR ORBITS FROM PLESETSK December 1996 Page 3

4 FIGURE 3 - PERFORMANCE CAPABILITIES IN ELLIPTICAL ORBITS FROM BAIKONUR FIGURE 4 - PERFORMANCE CAPABILITIES IN ELLIPTICAL ORBITS FROM PLESETSK 3.2 Spacecraft orientation and separation Thermal control manœuvres : Nominal payload separation velocity : 1 m/s Rotation rate : 0 rpmr Deployment mechanism type : spring release December 1996 Page 4

5 3.3 Payload interfaces Payload compartments and adaptors Payload fairing description - A basic m long fairing made of two aluminium shells has been used for protection of PROGRESS spacecraft (see Figure 5) but at least three other fairings are also available as shown in the following table. Length (m) External diameter (m) Mass (kg) (estimation) Mission PROGRESS spacecraft RESURS-F, FOTON satellites COSMOS type satellites Manned - A new serial fairing consists of two half-shell carbon fibre structure with a longitudinal type separation system (see Figure 6). Aluminium foil is applied to the internal and external surfaces of the nose fairing to protect against static electricity and to provide for optimal thermal conditions. A thermal insulation is applied to the forward cone external surface. The conical portion of the nose fairing is made of fiberglass honeycomb plastic. Payload access provision The standard fairing version includes two access doors per satellite. Four ports are provided to connect air ducts from the ground heat conditions maintaining air system. Payload adaptors interface The standard interface plane between the launch vehicle/orbital module and the payload is a bolted interface plane with a diameter of mm. For mounting a spacecraft on this mm interface, the user can either utilize one of the standard adaptors, in which case the spacecraft separation system is provided with the launch vehicle, or use any other type of adapter compatible with this interface, in which case the separation system has to be provided by the user. FIGURE 5 - STANDARD FAIRING (FOR PROGRESS SPACECRAFT) December 1996 Page 5

6 FIGURE 6 - PAYLOAD (AND ORBITAL MODULE) ARRANGEMENT UNDER NEW FAIRING 3.4 Environments Mechanical environment LEVEL (g) AT THE BASE OF THE SPACECRAFT Steady Longitudinal 4.3 state acceleration Lateral 0.4 NOTE Occurs at 1 st stage cut-off Low frequency vibration Longitudinal to 40 Hz Lateral to 10 Hz 10 to 20 Hz Random vibrations Longitudinal and lateral See Figure 7 December 1996 Page 6

7 (1) I stage function (2) II & III stage function (3) OM engine function FIGURE 7 - RANDOM VIBRATIONS DURING THE LAUNCH PHASE AND DURING ORBITAL MODULE (OM) ENGINE OPERATION Acoustic vibrations The total mean square acoustic pressure level (in the 20 to Hz frequency range) during flight does not exceed 142 db. Operation time comprises 60 s. Acoustic noise parameters under the fairing on the external S/C surface are shown in Figure 8. FIGURE 8 - MEAN SQUARE ACOUSTIC PRESSURE LEVELS AT THE SPACECRAFT EXTERNAL SURFACE FOR OCTAVE FREQUENCY SUBRANGE December 1996 Page 7

8 3.4.3 Shock FIGURE 9 - SHOCK SPECTRUM DURING LAUNCH VEHICLE SYSTEM OPERATION AND AT ORBITAL MODULE SEPARATION Thermal environment Prelaunch environment Temperature range inside fairing is adjustable between 15 C and 25 C (injected air); outlet temperature of air < 25 for spacecraft radiating less than 400 W. Cleanliness is class ; relative humidity 1% and filtration 0.2 µm. In-flight temperature under fairing The thermal flux radiated by the fairing does not exceed 800 W/m 2 at any point. Minimal temperatures are presented in Figure 10. Aerothermal flux after fairing jettisoning The thermal flux density radiated after fairing jettisoning at a 1 m radius ring surface oriented normally to the velocity vector under typical launch operation varies from 10 kw/m 2 to zero within 50 s after fairing jettisoning. December 1996 Page 8

9 FIGURE 10 - THERMAL ENVIRONMENT UNDER THE FAIRING Variation of static pressure under fairing 3.5 Operation constraints Ground constraints FIGURE 11 - VARIATION OF STATIC PRESSURE UNDER THE FAIRING Coordination is exercised by RKA Space Agency and the Ts SKB General Designer representing the entire launch authority. The "safety regulations" define the rules applicable to all operations including the use of hazardous systems or products. Launch rate capability 15 to 25 per year (nominal) (up to 45 launches in 1978 and 1979 but decreasing to 15 in 1994) Procurement lead time About 18-month cycle between contract signature and launch December 1996 Page 9

10 4. LAUNCH INFORMATION 4.1 Launch site Location Two sites are used: - the Baikonur cosmodrome (45 60' N, 63 40' E) near Tyuratam in Kazakhstan (2100 km to the South- East of Moscow). At least two operational launch pads are available. A layout of the launch complex is presented in Figure the Plesetsk cosmodrome (62 80' N, 40 30' E), a military site in the North-West Russia at some 800 km of Moscow, since 1993; the launching rate is lower than in Baikonur. Currently, /MOLNIYA launch vehicles are supported by three active pads (Complexes 16 and 43, left and right; see Figure 13), while a fourth pad (Complex 41) is in mothballs. Payload processing After arrival at the cosmodrome, the payload undergoes checkout and fueling. It is then transported to the Space Vehicle Assembly Building (MIK) to be integrated with the launch vehicle, the components of which have already been transported to the cosmodrome by rail. The integration process can follow one of two approaches, depending on payload requirements. In one option, the payload, adaptor, interstage are integrated vertically. The two halves of the payload fairing are then attached to the interstage () and connected together. This assembly (referred to as the "head block" in Russian terminology) is then rotated to an horizontal position, moved to a transporter/erector railroad car, and integrated horizontally with the rest of the launch vehicle. In the other option, the payload, interstage and core stage 2 are integrated vertically, after which the assembly is rotated to an horizontal position and the complete payload fairing is rolled into place over the payload. This stack is then transferred to a transporter/erector car and attached to the previously assembled core stage 1 and strap-ons. Throughout this process, temperatures within the MIK are maintained between 5 and 35 C. Once the payload is integrated with the launch vehicle, temperatures within the payload fairing can be maintained between 15 and 25 C. Launch vehicle processing Three or four weeks before launch, the components of launch vehicle are delivered to the MIK assembly complex in seven parts (4 strap-on boosters, 1 st stage, 2 nd stage and 3 rd stage). In few days, the separate parts are horizontally mated. After successful integration test, the entire launch vehicle with its payload is carried on railway transport erector car to the launch pad, tilted up, to sit on a stand over a large flame deflector pit. December 1996 Page 10

11 FIGURE 12 - BAIKONUR LAUNCH COMPLEX LAYOUT FIGURE 13 - PLESETSK COSMODROME December 1996 Page 11

12 The launch structure for the employs a launch system with four releasable support beams which accept the weight of the vehicle. The launcher is suspended over the gas deflecting trough with its tail portion 7 m below the level of the platform. After ignition, when slowly begins to rise, the four support arms, initially held in place by the weight of the launch vehicle, fall back under the influence of their large counter-weights and open like petals to clear the climbing vehicle. The entire processing flow, from launch vehicle and payload arrival at the cosmodrome until launch, requires 121 h, including 18 hours for on-pad activities. 4.2 Sequence of flight events A typical flight sequence for LEO mission is given hereunder. FLIGHT TIME EVENTS Lift-off; strap-on boosters and stage 1 ignite simultaneously Strap-on boosters separation Payload fairing jettison Core stage 1 separation; stage 2 ignition Stage 2 shutdown; spacecraft orbital injection Nota: the flight time, given here as an example, depends on the mission. December 2002 Page 12

13 4.3 Launch record data The information on this page is based on the TRW Space Log and launch history data provided by the Central Specialized Design Bureau (CSDB). Many discrepancies exist between these two sources for the years prior to The table below includes both VOSKHOD and launches. LAUNCH YEAR BA PL NUMBER OF SUCCESSFUL LAUNCHES NUMBER OF FAILURES TOTAL December 2004 Page 13

14 Failures 36 failures have been reported. Until 1990, no detail has been provided. In 1996, two successive launches failed. The last failure occured in 2002 and ended a long serie of successful flights. LAUNCH DATE RESULT CAUSE COSMOS At T + 49 s, the payload fairing broke in four pieces. The launch vehicle was destroyed at T s. The glue between layers of the glass reinforced plastic fairing had been incorrectly applied. The latches were also defective and unable to tolerate the aerodynamic pressure COSMOS At T + 50 s, the payload fairing desintegrated. It was a near repeat of the precedent launch failure. Same cause FOTON M1 At T + 29 s, the launch vehicle exploded. One of the RD-107 engine pressure fell down because of a leak in the nitrogen gas generator fed line. Previsional reliability: - Success ratio (table): 1 103/1 139 = 96.8% For the period the success ratio is 98.5%. 4.4 Planned launches Up to 12 launches are planned in December 2004 Page 14

15 5. DESCRIPTION 5.1 Launch vehicle FIGURE 14 - U (MANNED VERSION) LAUNCH VEHICLE 5.2 Overall vehicle Overall length Maximum diameter Lift-off mass (approx.) : 43.5 m (50.7 m for manned version with escape tower) : 10.3 m : 304 t (with 5.7 t payload) to 310 t December 2002 Page 15

16 5.3 General characteristics of the stages STAGE Designation Blocks B, V, G, D Block A Block I Manufacturer Ts SKB Ts SKB Ts SKB Length (m) Diameter (m) 2.68 (at base) Dry mass (t) x Propellant: Type Liquid Liquid Liquid Fuel Kerosene Kerosene Kerosene Oxidizer LO 2 LO 2 LO 2 Propellant mass (t) Fuel Oxidizer TOTAL x Tank pressure (bar) Total lift-off mass (t) Upper part DESIGNATION Manufacturer Mass (t) - Launch vehicle growth VEHICLE EQUIPMENT BAY FAIRING In order to enable to launch cluster satellites, it is planned to mount an upper stage designated IKAR beneath the fairing. It has also been envisioned to introduce a versatile family of 3 uprated launch vehicles, called RUS. A first model called 2 will consist of a "standardised upgraded package" with an available additional FREGAT upper stage for higher orbits. - Ballistic phase capability : - December 1996 Page 16

17 5.4 Propulsion STAGE Designation RD-107 RD-108 RD-0110 Engine Manufacturer ENERGOMASH ENERGOMASH KHIMAUTOMATIKI Number of engines 4 (4 chambers + 2 verniers each) 1 (4 chambers + 4 verniers) 4 (4 chambers + 4 verniers) Engine mass (t) Feed syst. type Turbopump Turbopump Turbopump Mixture ratio Chamber pressure (bar) Cooling Liquid (kerosene) Liquid (kerosene) Specific impulse (s) Sea level Vacuum Thrust (kn) Sea level x Vacuum x Burning time (s) Nozzle expansion ratio Restart capability No No No 5.5 Guidance and control Guidance Inertial Two avionics systems are used. One is located on stage 1 and controls stage 0 and stage 1. A second independent avionics package on stage 2 controls this stage Control STAGE Pitch, yaw, roll Precision Movable aerodynamic fins and 8 gimballed verniers 4 gimballed verniers 4 gimballed verniers December 1996 Page 17

18 6. DATA SOURCE REFERENCES 1 - World Space Systems Briefing - ELV - Teal Group - September International Reference Guide to Space Launch Systems - S.J. ISAKOWITZ - 2 nd edition AIAA Jane's Space Directory Europe & Asia in Space Kaman Sciences Corporation and USAF Phillips Laboratory - Pages Launcher User's Manual - Version 01 - Ts SKB & Starsem - April Aviation Week , pages Flight International - 3/ , page 37-7/ , page Encyclopedia Astronautica (astronautix.com/lvs/soy1a511.htm) December 2002 Page 18