SDO YUZHNOYE’S CAPABILITIES IN SPACE DOMAIN
INTERNATIONAL EU-RUSSIA/CIS CONFERENCE ON TECHNOLOGIES OF THE FUTURE: SPAIN-ISTC/STCU COOPERATION MADRID, APRIL 22-23, 2010
LAUNCH SERVICES ZENIT-3 SL
Zenit-3SL LV presents an optimal solution in terms of power characteristics, reliability, accuracy and cost of SC injection, which has been reached owing to utilization of developed systems and optimal planning of production process, transportation technology, prelaunch preparation and launch.
Zenit-3SL ILV is designed in accordance with monoblock tandem scheme and contains:
- the first and second stages of Yuzhnoye’s development; - upper stage elaborated by Energia RSC; - payload unit elaborated by Boeing company.
Zenit-3SL is operated under Sea Launch Program and launched from Odysseus floating platform in the equatorial area of the Pacific Ocean MAIN CHARACTERISTICS OF ZENIT-3 SL
Maximum lift-off mass, t 473 Payload mass injected into the GSO 6,0 transfer orbit, t: Propellant mass, t 425 Engines thrust during LV start, tf 740 Maximum acceleration during 4,0 injection, g Propellants liquid oxygen/kerosene Full length, m 59,6 Diameter of LV stages (1st and 2nd), m 3,9 Upper stage diameter, m 3,7 Nose fairing diameter, m 4,15 Number of stages 3 Service beginning March1999 ZENIT-2 SLB and ZENIT-3 SLB
Zenit-2 SLB and Zenit-3 SLB are designed for launching under the LAND LAUNCH program from launch site in Baikonur MAIN CHARACTERISTICS OF ZENIT-2 SLB and ZENIT-3 SLB
Integrated launch vehicle Zenit-3SLB ILV Zenit-2SLB ILV Maximum lift-off mass, t 466.2 458.2 Payload mass injected into orbits, t:
ISS: Нcirc=400 km, i=51,6° - 12.03 GSO transfer: H=35786x230 km, i=0° 3.75 - Propellant mass, t ` 410 Engines thrust during LV 740 start, tf
Maximum acceleration 4.0 4.0-6.0* during injection, g
Propellants liquid oxygen/kerosene Full length, m 58.65 57.35 Diameter of LV stages 3.9 3.9 (1st and 2nd), m
Upper stage diameter, m 3.7 - Nose fairing diameter, m 4.1 3.9 Number of stages 3 2 DNEPR
Distinctive Features
. Deployment of satellites of 3500…300 kg into circular orbits with the heights of 300…900 km . Commercial operation was commenced by deployment of SSTL (Great Britain), Italian, Saudi Arabian and Malaysian satellites (1999-2000) . High accuracy of deployment . Low cost . High reliability . Flexibility MAIN CHARACTERISTICS OF DNEPR
Lift-off mass (at SC of 2000 kg), kg: first stage 208900 second stage 47380 third stage 6266 Propulsion components oxidizer amil fuel geptil Propellant mass: first stage 147900 second stage 36740 third stage (main mode/throttled mode) 1910 Vacuum thrust, тf first stage 461,2 second stage 77,5 third stage 1,9/0,8 Flight reliability 0,97 SC injection accuracy for Hcirc=300 km: orbit altitude, km 4,0 revolution period, s 3,0 inclination, ang. min 0,04 right ascension of ascending angle, deg 0,05 Inclinations of orbits 50,5 ; 64,5 ; 87,3 ; 98 CYCLONE-4
In the near future Yuzhnoye will be able to propose newly developed launch vehicle for providing customers with reliable launch services. The new system is Cyclone-4 launch vehicle (launch services will be provided by Alcantara Cyclone Space Joint Stock Company). Currently Yuzhnoye’s experts are working on increasing of payload capabilities of Cyclone-4 LV
Launch site: Alcantara (Brazil)
Launch site has convenient geographical position and provides wide range of launch azimuths Payload capability for GTO is: 1600 kg (Alcantara, i=5.1 degrees) CYCLONE-4
Number of stages 3 Lift-off mass, t (without PL) 191 Propellant components PAYLOAD CAPABILITIES Oxidizer NTO Fuel UDMH Propellant mass, t 6000 1st Stage 121 5000 2nd Stage 49 4000 3rd Stage 9 Vacuum thrust 3000 of engines, tf Gpl, kg 2000 1st Stage 303 1000 2nd Stage 101.5 0 3rd Stage 7.9 2000 3000 4400 6000 8000 Injection accuracy Hcirc, km for Hcirc=500 km, i=90º: On altitude, km ±5 On inclination, deg ±0.05…0.08 for GTO: On altitude, km 5.0 (perigee) 100 (apogee) On inclination, deg ±0.05…0.08 SPACECRAFT
MS-2-8 OPTOELECTRONIC EARTH OBSERVATION SATELLITE
MS-2-8 satellite is intended for acquisition of the digital images of Earth surface in panchromatic and multi-spectral bands with resolution of <8 meters, and middle infrared spectral band with resolution of ~46 meters. The satellite consists of the platform and payload. Egyptsat-1 remote sensing satellite developed by Yuzhnoye was successfully launched on April 17, 2007 by Dnepr LV from Baikonur launch site MS-2-3 OPTOELECTRONIC EARTH OBSERVATION SATELLITE
MS-2-3 satellite is intended for acquisition of the digital images of Earth surface in panchromatic band with resolution of <3 meters, and multi-spectral band with resolution of <8 meters. Satellite is designed using the MS-2-8 satellite platform. S-1-2 OPTOELECTRONIC EARTH OBSERVATION SATELLITE
S-1-2 satellite is intended for acquisition of the digital images of Earth surface in panchromatic band with resolution of <2 meters, and multi- spectral band with resolution of <6 meters. The satellite consists of the platform and payload. S-3-O OPTOELECTRONIC EARTH OBSERVATION SATELLITE
S-3-O satellite is intended for acquisition of the digital images of Earth surface in panchromatic band with resolution of <1 meters, and multi-spectral band with resolution of <3 meters. The satellite consists of the platform and payload. MS-2-RL RADAR EARTH OBSERVATION SATELLITE
MS-2-RL satellite is intended for acquisition of the radar images of Earth surface with resolution of ~2000 meters or ~200 meters. Satellite is designed using the MS-2-8 satellite platform. ROCKET ENGINES More than 35 liquid rocket engines and liquid propulsion systems have been developed
The achieved level of reliability
For Liquid Rocket Engines For Solid propellant no less than Rocket Motors 0,992-0,999 0,995-0,999 MAIN SPECIFICATIONS OF THE ENGINES RD858 AND RD859
Propellants NTO+UDMH Vacuum thrust, kgf – main engine from 400 to 2250 – backup engine 2045 Vacuum specific impulse, s – main engine 315 – backup engine 312 Burn time, s – main engine up to 470 – backup engine up to 400 Engine cluster mass, kg 110 Mixture ratio – main engine from 1,6 to 2,03 – backup engine 2,0 MAIN ENGINE ASSEMBLY FOR THE EUROPEAN VEGA LV
The Main Engine Assembly is a part of Liquid Propulsion System for Attitude Vernier Upper Module of the European Vega Launch Vehicle.
The MEA is developed under a Contract with Avio (Italy) on the basis of units from serially produced engines.
Purposes of the MEA are as follows: Thrust generation; Pitch/yaw control; Upper Module maneuvering; Upper Module deorbit.
Propellants NTO+UDMH Vacuum thrust, kgf 250 Vacuum specific impulse, s 315.5 Number of burns 5 ENGINE UNIT OF DU802 PROPULSION SYSTEM
The engine unit is a component of DU 802 liquid propulsion system
Propellants NTO+UDMH
Vacuum thrust, kgf 450 Vacuum specific impulse, s 322.5 Mixture ratio 2.25 Number of burns up to 10 Burn duration of a single run, s: - max 350 - min 3 Total burn duration, s 350
A phase of development tests has been accomplished DU802 PROPULSION SYSTEM Currently the Upper Stage for Ukrainian – Russian Dnepr LV has been designed and is now under intensive testing. A principally new propellants supply system was introduced with the use of Pneumopump Assembly (PPA). Utilization of PPA allows to increase power-mass characteristics, some of which cannot be reached by the existing propulsion systems.
Propellants: • Oxidizer NTO • Fuel UDMH Propellants mass, kg 250–500 Dry mass, kg 165.4
Vacuum specific impulse, s - Main engines (ME) 322,5 - Thruster 243 Vacuum thrust, kgf - ME 450 - Thruster 11,1 Mixture rate 2,25 Number of ME burns 10
Propellants supply system has been tested (pressurization system, propellants tanks and PPA). LIQUID ROCKET ENGINE RD861K The engine RD861K is designed for generating thrust and ensuring control in pitch and yaw channels through an active leg of the Cyclone-4 LV third stage flight.
Propellants: - Oxidizer NTO - Fuel UDMH Vacuum thrust, kgf 7916 Vacuum specific impulse, s 330 Mass, kg 207 Number of burns 3…5 Total burn duration, s 450 Length, mm 2000 Diameter of nozzle exit, mm 1010 PROJECT OF THE ENGINE WITH THRUST VALUE 120T (RD801)
Engine Performances: • propellant components RG-1+ O2 • thrust, tf: • vacuum 133.55 • earth 120 • mixture ratio 2.6 • specific impulse, sec: • vacuum 335 • earth 300 • engine mass (dry), kg 150030 PROJECT OF THE ENGINE WITH THRUST VALUE 200T
Engine Performances: • propellant components RG-1+ O2 • thrust, tf: • vacuum 203,9 • earth 182 • mixture ratio 2.65 • specific impulse, sec: • vacuum 335 • earth 299,1 • engine mass (dry), kg 2720100 MAIN ENGINES FOR LV UPPER STAGES ON THE BASIS OF STEERING ENGINE RD-8
RD-8 Thrust in vacuum 8 tf.
Four-chamber main engine RD809. One-chamber main Thrust in vacuum 9 tf. One-chamber main engine RD809К. engine RD802. Thrust in vacuum Thrust in vacuum 2 tf. 8…10 tf. PROSPECTIVE LAUNCH TECHNOLOGIES MAYAK ROCKET SPACE COMPLEX MAYAK LAUNCH VEHICLES FAMILY
MAYAK-AB-C
Diameter # of stages # of boosters
1 - 3m 2 - 3,9m 3 - <3,0m
Launch mass, t 8090 150180 200230 310360 341060
# of stages/boosters 2/02/0 2/02/0 2/42/4 2/2/00 2/42/4
Payload mass, kg
H=200 km, i=50O 10001500 3000 60006000 810000000 1000012000 H=150/35800H=150/35790 km, i=1i=16,8O ° - - -500 15001700 23300000 30003700 MICROSPACE
Purpose of MICROSPACE is the injection of microsatellites into a wide orbit range using supersonic aircraft LAUNCH VEHICLE
Characteristics 1st stage 2nd stage 3rd stage
Dry weight of separable part of the stage, kgf 621 272 110
Propellant weight, kgf 3925 770 490 Propellant type НТРВ НТРВ НТРВ
Average thrust level in vacuum, kgf 16395 3166 3548
Vacuum specific impulse of engine thrust, kgf s/kg 289.4 287.2 292.6
Launch weight of LV is ~6300 kgf, including SC – 40 kgf
TWO-STAGE LV
1800
17600 LV Main Specifications
Performances I stage II stage
Propellant НТРВ NТ+ UDMH Propellant weight, kgf 24104 8500 Fairing weight, kgf - 190 Weight of stages (without fairing weight and spacecraft), kgf 36338 9659 Average vacuum thrust, kgf ~126650/52700 ~7900 Specific vacuum thrust, kgf s/ kg 280 330 Number of engines ignitions 1 Up to 3 times Gas load, kgf - ~10
Space rocket lift-off weight ~ 37030kgf, including SC weight = 500kgf, injected into orbit Нcir=500km, i=0
THREE-STAGE LV
1800
~ 26000
LV Main Specifications
Performances 1 stage II stage III stage
Propellant НТРВ HTPB NТ+ UDMH Propellant weight, kgf 24400 24104 8500 Fairing weight, kgf - - 315 Weight of stages (without fairing weight and spacecraft), kgf 63235 36357 9659 Average vacuum thrust, kgf 123800 126650/52700 7916 Specific vacuum thrust, kgf s/ kg 274 280 330 Number of engines ignitions 1 1 Up to 3 times Gas load, kgf - - ~10
Space rocket lift-off weight ~ 64550kgf, including SC weight = 1000kgf, injected into orbit Нcir=500km, i=0 GLOBAL SPACE PROJECTS SOLAR KEY PROJECT
Re-reflecting spacecraft Re reflected energy ray
Focal line
Reflecting spacecraft Solar radiation density Energy ray receiver area
Earth shadow
Re-reflecting spacecraft
550 thousand km. Working part of concentration area
1200 thousand km. SPACE DISPOSAL OF HAZARDOUS WASTE
The analysis of the ecological situation involving the increasing amount of waste of the atomic power stations shows that the issue of the waste isolating becomes extremely urgent for mankind. Isolation in space shall enable the globe to get rid of the long-lasting radioactive wastes forever unlike any other ways of burial on the planet. COMMERCIAL AIRCRAFT PROTECTION SPATIAL DISPLACEMENT OF THERMAL IMAGE SDTI technology is a novel countermeasure to the threat posed by Man- Portable Aircraft Defense Systems (MANPADS), which are typically heat- seeking surface-to-air guided missiles targeted at aircraft by individual/ terrorists in the vicinity of airports.
Unique countermeasure system is based on rocket engine technologies and generates the signals which jam the infrared guidance systems of the missile with high probability.
The system provides a highly innovative concept for civilian aircraft protection from potential terrorist attacks by heat-seeking missiles during the whole phase of mission: take-off, flight and landing. SDTI ADVANTAGES
Low cost – SDTI is expected to be less than half the cost of the countermeasure technologies currently being tested and evaluated for commercial use.
No missile warning system required – SDTI is a passive countermeasure; it operates continuously during ascent and descent (while the aircraft is vulnerable to MANPADS missiles). The system requires no missile detection and tracking sensor, and hence generates no false alarms.
SDTI is equally effective against multiple simultaneous missile launches. Other directed countermeasures can only defeat one missile at a time.
Airline friendly – SDTI does not include any classified hardware or software and requires no special operator training.
Low power requirement – the Gas Dynamic Thermal Generator is a very efficient source of broadband IR energy. THANK YOU! CONTACT INFORMATION:
Diana Kolova Senior Manager, Business Development Yuzhnoye State Design Office
E-mail: [email protected]
Tel: +38 056 770 04 47 Fax: +38 056 770 01 25