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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
Maximum lift-off mass, t 473
Payload mass injected into the GSO transfer orbit, t:
6,0
Propellant mass, t 425
Engines thrust during LV start, tf 740
Maximum acceleration during
injection, g
4,0
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
MAIN CHARACTERISTICS OF ZENIT-3 SL
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
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°
GSO transfer:
H=35786x230 km, i=0°
-
3.75
12.03
-
Propellant mass, t ` 410
Engines thrust during LV start, tf
740
Maximum acceleration during injection, g
4.0 4.0-6.0*
Propellants liquid oxygen/kerosene
Full length, m 58.65 57.35
Diameter of LV stages (1st and 2nd), m
3.9 3.9
Upper stage diameter, m 3.7 -
Nose fairing diameter, m 4.1 3.9
Number of stages 3 2
MAIN CHARACTERISTICS OF
ZENIT-2 SLB and ZENIT-3 SLB
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
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
MAIN CHARACTERISTICS OF
DNEPR
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
CYCLONE-4
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)
Lift-off mass, t (without PL)
9
7.9
101.5
±5
±0.05…0.08
303
3
191
UDMH
49
121
NTO
Number of stages
Propellant components
Oxidizer
Fuel
Propellant mass, t
1st Stage
2nd Stage
3rd Stage
Vacuum thrust
1st Stage
2nd Stage
3rd Stage
Injection accuracy
On altitude, km
On inclination, deg
for Hcirc=500 km, i=90º:
of engines, tf
5.0 (perigee)On altitude, km
for GTO:
±0.05…0.08On inclination, deg
100 (apogee)
0
1000
2000
3000
4000
5000
6000
2000 3000 4400 6000 8000
Hcirc, km
Gp
l, k
g
PAYLOAD CAPABILITIES
CYCLONE-4
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
no less than
0,992-0,999
For Solid propellant
Rocket Motors
0,995-0,999
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 SPECIFICATIONS OF
THE ENGINES RD858 AND
RD859
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
MAIN ENGINE ASSEMBLY FOR
THE EUROPEAN VEGA LV
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
ENGINE UNIT OF DU802
PROPULSION SYSTEM
A phase of development tests has been accomplished
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
DU802 PROPULSION SYSTEM
Propellants supply system has been
tested (pressurization system,
propellants tanks and PPA).
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
LIQUID ROCKET ENGINE
RD861K
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 120T (RD801)
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
Four-chamber main engine RD809.
Thrust in vacuum 9 tf.One-chamber main
engine RD809К. Thrust in vacuum
8…10 tf.
One-chamber main engine RD802.
Thrust in vacuum 2 tf.
RD-8Thrust in vacuum 8 tf.
MAIN ENGINES FOR LV UPPER STAGES
ON THE BASIS OF STEERING ENGINE RD-8
PROSPECTIVE LAUNCH
TECHNOLOGIES
MAYAK ROCKET SPACE COMPLEX
MAYAK LAUNCH VEHICLES FAMILY
Launch mass, t
# of stages/boosters
H=200 km, i=50O
H=150/35790 km, i=1O
Payload mass, kg
180
2/0
500
230
2/4
60001700
360
2/0
100003000
410
2/4
120003700
Diameter
1 - 3m 2 - 3,9m 3 - <3,0m
# of stages # of boosters
MAYAK-AB-C
90
2/0
1500-H=150/35800 km, i=16,8°
80
2/0
1000
-
150
2/0
3000
-
200
2/4
6000
1500
310
2/0
8000
2300
360
2/4
10000
3000
Purpose of MICROSPACE is the injection of microsatellites into a wide
orbit range using supersonic aircraft
MICROSPACE
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
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
LV Main Specifications
17600
1800
THREE-STAGE LV
~ 26000
1800
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
LV Main Specifications
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
Focal line
Re reflected energy ray
Energy ray receiver
550 thousand km.
1200 thousand km.
Working part of concentration area
Earth shadow
Re-reflecting spacecraft
Reflecting
spacecraft
Solar radiation density
area
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
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.
SPATIAL DISPLACEMENT OF THERMAL IMAGE
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: space@yuzhnoye.com
Tel: +38 056 770 04 47
Fax: +38 056 770 01 25
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