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