CIS
PROTON M
1. IDENTIFICATION
1.1
Name
PROTON M
1.2
Classification
¾ Family
¾ Series
¾ Version
:
:
:
PROTON
¾ Category
¾ Class
¾ Type
M
1.3
Manufacturer
:
KHRUNICHEV Entreprise
Novoza Vodskaya ulitsa, 18
MOSCOW
121309 Russian Federation
Fax: (095) 142-59-00
1.4
Development manager :
KHRUNICHEV SRPSC
Novoza Vodskaya ulitsa, 18
MOSCOW
121309 Russian Federation
1.5
Vehicle operator
1.6
Launch service agency :
:
:
:
SPACE LAUNCH VEHICLE
Heavy Launch Vehicle (HLV)
Expendable Launch Vehicle (ELV)
:
International Launch Services (ILS) Through Lockheed - Khrunichev Energia International (LKEI)
a US - Russian joint venture
SAN DIEGO, CALIFORNIA, USA
MOSCOW, RUSSIA
1.7
Launch cost
:
About 85 M$
2. STATUS
2.1
Vehicle status
:
Operational
2.2
Development period
:
1992-1999
2.3
First launch
:
07.04.2001
December 2001
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PROTON M
3. PAYLOAD CAPABILITY AND CONSTRAINTS
3.1
Payload capability
3.1.1 Low Earth Orbits
ORBIT TYPE
LEO CIRCULAR
LEO CIRCULAR
LEO ELLIPTICAL
Altitude
(km)
(Perigee/Apogee)
175
170
186 x 222
Inclination
64.8
72.7
51.6
BAIKONUR
BAIKONUR
BAIKONUR
20 610
19 975
21 000
(°)
Site
Payload mass (kg)
3.1.2 Geosynchronous and Interplanetary Orbits
ORBIT TYPE
Altitude
(km)
(Perigee/Apogee)
Inclination
GTO
GEO
Lunar
Mars transfer
5 500 x 35 786
35 786
-
-
25
0
BAIKONUR
BAIKONUR
BAIKONUR
BAIKONUR
5 200-5 500(1)
2 920
5 600
4 800
(°)
Site
Payload mass (kg)
(1) See Figure 1
3.1.3 Injection accuracy (3 sigma)
PERIGEE
APOGEE
INCLINATION
PERIOD
200 km circular support orbit
± 6 km
± 15 km
± 0.025°
±8s
1 000 km circular orbit
± 10 km
± 10 km
± 0.05°
± 100 s
5 500 km x 36 000 km
(i: 25 deg) GTO
± 400 km
± 150 km
± 0.5°
± 550 s
GEO
December 1999
ECCENTRICITY
LONGITUDE
INCLINATION
PERIOD
0.009
± 1°
0.75°
± 20 min
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PROTON M
Figure 1 plots optimum GTO playload versus DELTA Velocity to GEO for three performance variants of
PROTON / BREEZE M.
-
Configuration 1: performance capability for PROTON K/BREEZE M (initial flights n° 1, 2, 3) and
contrained performance for PROTON M/BREEZE M (maiden flight n° 4); a 4 800 kg payload mass can
be delivered to a reference GTO in this case.
-
Configuration 2: performance for PROTON M/BREEZE M introduced during second half of 2000
(flights n° 5, 6, 7) (5 200 kg)
-
Configuration 3: performance for mature flight configuration of PROTON M/BREEZE M (flights n° 8
and on) (5 500 kg).
FIGURE 1 - PROTON M/BREEZE M PERFORMANCE TO REPRESENTATIVE GEOSYNCHRONOUS
TRANSFER ORBITS
3.2
¾
¾
¾
¾
Spacecraft orientation and separation
Thermal control manœuvres
Nominal payload separation velocity
Rotation rate
Deployment mechanism type
December 1999
: yes
: ≥ 0.3 m/s
: up to 9°/s
: spring release
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3.3
PROTON M
Payload interfaces
3.3.1 Payload compartments and adaptors
¾ Payload fairing description
The PROTON M has at least two payload fairings available with the BREEZE M upper stage: a standard
commercial fairing with a usable volume comparable to that of the Block DM, and a long version of the
fairing. Each of these fairings is assembled from two half-shell structures. The structure consists of
aluminum skin, stringer, and frame construction. The geometric characteristics of the PROTON M fairings
are shown in Figure 2. These fairings typically enclose both the payload and any supplemental orbital
propulsion units employed by the payload.
FIGURE 2 - BREEZE M PAYLOAD FAIRINGS (STANDARD AND LONG)
Main Dimensions:
¾ Overall length
¾ Diameter
¾ Volume
: 11.60 m (standard), 13.20 m (long)
: 4.35 m (to be confirmed)
: 100 m3 (standard)
¾ Payload adaptor interface
The Breeze core structure provides the payload adaptor and electrical interfaces to the customer's
spacecraft. The interface between the stage and its payload adapter is 2 490 mm in diameter, allowing
the Breeze M to accommodate large diameter payload adaptors and a static bending moment about this
interface of 18,000 kg-m. Any spacecraft produced with adaptor interfaces compatible with one of the
defacto industry standards should be able to be accommodated on the PROTON M without modification
of these interfaces. The early selection of an appropriate adaptor should be coordinated with ILS.
December 1999
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PROTON M
¾ Payload access provisions
Four fairing access doors are located on the lower part of the fairing structure. These doors are nominally
used for access to the BREEZE M. The customer may use these doors for access to spacecraft related
interface equipment. Up to 2 access doors may be provided in the locations shown in Figure 3.
VIEW A-A
VIEW B-B
FIGURE 3 - FAIRING GENERAL LAYOUT (ACCESS LOCATIONS)
3.4
Environments
3.4.1 Mechanical environment
This table provides the BREEZE M quasi-static loads in longitudinal and lateral axes at the spacecraft center
of mass.
EVENT
Lift-off
Maximum Dynamic Pressure (Qmax)
LONGITUDINAL, g
LATERAL, g
2.3
1.35
-1.35
0.3
1.35
-1.35
2.2
1.2
-1.2
4.3
0.9
-0.9
1 /2
st
nd
stages before separation
1 /2
st
nd
stages after separation (max)
3
0.9
-0.9
1 /2
st
nd
stages after separation (max tension)
-3
0.9
-0.9
nd
rd
3
0.3
-0.3
rd
th
2.8
0.3
-0.3
2 /3 stages separation
3 /4 stages separation
December 1999
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PROTON M
Static and dynamic accelerations maximum 3 σ values at the spacecraft interface are shown in the following
table.
EVENT
LONGITUDINAL, g
Lift-off
1.5
Wind and Blast
(Qmax)
2.2
LONGITUDINAL
DYNAMIC, g
1.5
- 1.5
TRANSVERSE
DYNAMIC, g
1.1
- 1.1
0.5
- 0.5
st
3.6
0.9
- 0.9
0.9
- 0.9
After 1 stage
booster separation
1
2
- 2.8
0.9
- 0.9
2nd stage engine
cut-off
2.7
0.3
- 0.3
0.3
- 0.3
3rd stage engine
cut-off
2.8
0.3
- 0.3
0.5
- 0.5
st
Separation 1 /2
stages
nd
Before 1 stage
booster separation
st
3.4.2 Acoustic vibrations
The peak acoustic loads do not act longer than 3 s at lift-off and 50 s while passing through the zone of
maximum aerodynamic drag.
Acoustic load characteristics normalized to the threshold pressure of 20 µPa are shown in Figure 4.
FIGURE 4 - MAX EXPECTED ACOUSTIC ENVIRONMENT (THIRD OCTAVE)
December 1999
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PROTON M
3.4.3 Shock
Worst case shock levels are introduced into the spacecraft during the firing of the spacecraft/adapter
separation system. The level is dependent on the type of clampband and the clampband tension. For the
existing standard adapter configurations, three specific levels may be encountered as indicated in Figure 5.
The 1194 separation systems use either a 26.6 kN or a 40 kN preload and the shock levels differ
accordingly. The 1666 separation systems use a 30 kN preload and the corresponding shock level is as
indicated.
FIGURE 5 - PYROSHOCK SPECTRUM AT ADAPTER / PAYLOAD INTERFACE
3.4.4 Thermal environment
The on-pad air conditioning system remains active 24 hours a day until approximately 1.5 hours prior to
launch when preparations are begun for Mobile Service Tower rollback. To provide thermal conditioning of
the fairing after Mobile Service Tower rollback, a liquid thermal control system is provided in the fairing. This
system is known as the "LSTR" for Liquid System, Thermal Regulation. It consists of radiators on the fairing
inside wall connected to ethylene glycol filled pipes which run to a thermal control system in the launch pad
complex. This system is activated 3 hours prior to launch and purged with dry nitrogen 5 minutes prior to
launch to insure that the lines are free of liquid prior to lift-off. Should the launch be aborted, the liquid
system can be quickly reactived and the Mobile Service Tower will be brought up to renew air-conditioning
within 2 hours. A schematic of both the liquid and air thermal control systems is shown in Figure 6 along with
an approximate operational timeline.
During Ascent, the launch vehicle will be exposed to aerodynamic heating. Following fairing jettison, the
spacecraft will be exposed to solar radiation and free molecular heat flux. A thermal analysis will be
performed using the Customer supplied spacecraft thermal model to predict spacecraft temperatures during
this phase of the mission. The heat flux density radiated upon the spacecraft by the internal surfaces of the
²
internal of the fairing should not exceed 500 W/M from the time of launch until fairing jettisoned. For
commercial missions, the fairing is jettisoned at 342 - 344 seconds (121 - 125 km alt.) into flight and the free
²
molecular heat flux shall not exceed 1 135 W/M at any time following fairing jettison.
December 1999
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PROTON M
FIGURE 6 - FAIRING AIR AND LIQUID THERMAL CONTROL SYSTEM SCHEMATIC AND OPERATIONS
TIMELINE
¾ Pressure on fairing
During ascent, the payload compartment is vented through 4 venting holes distributed equally around the
cylindrical portion of the fairing. Maximum rate of pressure drop in the fairing will not exceed 3.5 kPa/s. A
representative pressure drop profile inside the fairing flight is given in Figure 7.
At the moment of fairing jettison, the pressure across the fairing halves shall not exceed 700 Pa.
December 1999
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PROTON M
FIGURE 7 – TYPICAL VENTING PROFILE DURING ASCENT
¾ Cleanliness level
The contamination environment around the spacecraft is controlled by use of class 100 000 clean room
facilities.
3.5
Operation constraints
¾ Ground constraints
: coordination is exercised by ILS and Khrunichev with SpaceCraft Customer
(SCC) organization.
¾ Launch rate capability : up to 12 per year.
¾ Procurement lead time : 24 months for non-recurring program
12 months for recurring program.
¾ Integration process
: a typical mission integration schedule is as shown in Figure 8.
FIGURE 8 - BASELINE INTEGRATION SCHEDULE
December 1999
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PROTON M
4. LAUNCH INFORMATION
4.1
Launch site (see also PROTON K data sheet)
The PROTON launch complex, consisting of spacecraft and launch vehicle processing and integration
facilities and four launch pads (two of which are available for commercial use), is located at the BAIKONUR
Cosmodrome. BAIKONUR, shown in Figure 9, is located approximately 2,000 km southeast of MOSCOW in
the Republic of KAZAKHSTAN. The BAIKONUR Cosmodrome measures approximately 90 km east-to-west,
and 75 km north-to-south, and supports many other launch vehicles, including the SOYUZ, VOSTOK, ZENIT
and ENERGIA. Temperatures range from - 40°C to 45°C during the year.
The PROTON launch system is designed to operate under the severe environmental conditions encountered
at BAIKONUR.
The PROTON can be launched year around, and the time between launches from an individual pad can be
as short as 25 days. PROTON has demonstrated a launch rate of four per month from multiple launch pads,
and a long term average launch rate of approximately twelve per year. The capability of the PROTON
system to launch in severe environmental conditions decreases launch delays and ensures that payloads
reach orbit as scheduled to begin revenue generating activities. The short turnaround time between launches
can ensure that spacecraft constellations are deployed quickly, minimizing the time required to enter service.
The basic launch structures are unchanged. The modified launch pad for PROTON M will also support
PROTON K/Block DM.
FIGURE 9 - BAIKONUR LAUNCH SITE (AVAILABLE DIRECT INJECTION INCLINATIONS)
December 1999
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PROTON M
¾ Launch vehicle processing
The PROTON LV stages and fairings are built in MOSCOW by KHRUNICHEV and transported by rail to the
BAIKONUR Cosmodrome typically well in advance of the SC delivery date. After transportation of the
PROTON's stages and fairing by rail, LV build-up takes place in an integration and test facility (Building 92-1)
capable of supporting four simultaneous PROTON assembly and checkout operations. The Fairing is moved
either to Building 40 or to Building 92A-50, depending upon which facility is to be used for SC integration,
prior to SC arrival, there it is stored and cleaned in preparation for encapsulation. The Fourth Stage is
delivered to Area 254 for pre-launch checkout and testing.The Fourth Stage is then delivered to the Building
44 in Area 31, the propellant fuelling hall, where MMH and N2 O4 are loaded. It is then moved to either
Building 40 or 92A-50 for integration with the spacecraft (see Figure 10). Payload adapters typically are
delivered shortly before the processing cycle and prepared in the Integration Hall (100A or 101) of whatever
processing area is being used for that program.
In advance of spacecraft arrival, Payload Processing Facilities undergo facility activation and certification.
Building 40, 40D, 44, or 92A-50 are verified to meet environmental control and cleanliness requirements, in
addition to commodities and power support requirements usually a week prior to the spacecraft arrival date.
FIGURE 10 - BAIKONUR FACILITIES MAP
December 1999
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4.2
PROTON M
Sequence of flight events
The first stages of the PROTON vehicle use a standard ascent trajectory to place the orbital unit (Breeze M
upper stage and payload) into a 200 km circular orbit inclined et 51.6°, 64.8° or 72.7°.
FIGURE 11 - TYPICAL PROTON BOOSTER ASCENT
FLIGHT TIME
- 1.60
December 1999
EVENTS
Stage ignition - 10% thrust
0.00
Begin stage 1 thrust to 100%
0.57
Lift-off
1.00
Stage 1 thrust to 100%
116.91
Stage 2 ignition
121.11
Stage 1/2 separation
330.00
Stage 3 vernier engine ignition
332.70
Stage 2 engine shutdown
333.40
Stage 2/3 separation
335.80
Stage 3 main engine ignition
344.20
Payload Fairing jettison
567.11
Stage 3 main engine shutdown
577.11
Stage 3 vernier engine shutdown
582.01
Stage 3 upper stage separation
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PROTON M
¾ Upper stage ascent
The Breeze/payload unit is placed into high-energy suborbital state by the third stage of PROTON. After
jettison of the third stage, the Breeze upper performs a small propulsive maneuver to deliver itself and the
attached satellite to a standard low earth parking orbit. After a coast of approximately 45 minutes, the
Breeze stage performs the second of four propulsive maneuvers. This second main burn is used to begin
the process of raising the apogee of the transfer orbit to geosynchronous altitude.
Figure 12 illustrates the main characteristics of the trajectory for a PROTON M launch to geosynchronous
transfer orbit.
FIGURE 12 - TYPICAL BREEZE M FLIGHT TO GTO
December 2004
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4.3
PROTON M
Launch record data
LAUNCH DATE
SITE
NUMBER OF
SATELLITES
ORBIT
RESULT
REMARK
07.04.01
BAIKONUR
1
GTO
Success
-
29.12.02
BAIKONUR
1
GTO
Success
-
15.03.04
BAIKONUR
1
GTO
Success
-
16.06.04
BAIKONUR
1
GTO
Success
-
04.08.04
BAIKONUR
1
GTO
Success
-
14.10.04
BAIKONUR
1
GTO
Success
-
¾ Failures
: none
¾ Previsional reliability : ¾ Success ratio
4.4
: 100% (6/6)
Planned launches
ILS is planning to launch six or seven PROTON M/ BREEZE M in 2005.
December 2004
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PROTON M
5. DESCRIPTION
PROTON M evolution versus PROTON K previous version
Compared to PROTON K, the main enhancements are:
¾ Stage 1 RD-253 engines thrust increased by 7% (this enhancement is accomplished with only a minor
modification to the propellant flow control valves)
¾ Mass reductions for stages 1, 2 and 3 by reducing cross sectional area structural components
¾ Propellant feed systems of stages 1 and 2 simplified and redesigned to reduce propellant residuals by
50%
¾ Reinforced design of the stage 2 front bay and stage 3 aft skirt
¾ New instrumentation bay with new digital control system
¾ Upper stage BREEZE M with improved performance. Sized to accommodate greater mass and volume.
Potential dual payload capability of Spelda type
¾ Payload fairing with larger useable payload envelope
5.1
Launch vehicle
FIGURE 13 - PROTON M
December 2004
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5.2
PROTON M
Overall vehicle
¾ Overall length
¾ Maximum diameter
¾ Lift-off mass (approx.)
5.3
: 57.2 m (with standard commercial fairing)
: 7.4 m
: 691.27 t
General characteristics of the stages
STAGE
1
2
3
Designation
4
Breeze M
Manufacturer
KRUNICHEV
KRUNICHEV
KRUNICHEV
KRUNICHEV
Length (m)
21.18
17.05
4.11
2.61
Diameter (m)
7.40
4.10
4.10
4.10
Dry mass (t)
30.60
11.40
3.70
2.37
¾ Type
Liquid
Liquid
Liquid
Liquid
¾ Fuel
UDMH
UDMH
UDMH
UDMH
N 2O 4
N 2O 4
N 2O 4
N 2O 4
419 410
156 110
46 560
19 800
Propellant:
¾ Oxidizer
Propellant mass (kg)
¾ Fuel
¾ Oxidizer
¾ Water
TOTAL
Tank pressure
(bar)
Total lift-off
mass (t)
¾ Upper part
DESIGNATION
Manufacturer
VEHICLE EQUIPMENT
BAY
FAIRING
NPO Mars and NII AP
(guidance systems)
Mass (t)
¾ Launch vehicle growth
Use of a cryogenic stage 4 called KVRB (KB Khimmach engine).
The first flight tests of this stage are planned for 2003.
December 2004
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CIS
5.4
PROTON M
Propulsion
STAGE
Designation
Manufacturer
2
3
4
RD-253
RD-0210
RD-0210
11 DM 58
MPO Energomash
Number of engines
4
1 300
566
Turpopump
Turpopump
Mixture ratio
2.69
Chamber pressure
(bar)
158
Cooling
KB Khimautomatiki KB Khimautomatiki
6
Engine mass (kg)
Feed syst. type
1
1 (+ 4 verniers)
KB Khimmach
1
95
Turpopump
Turpopump
2.00
150
150
326.5
326.5
325.5
2 320
583 (+ 31)
19.6
up to 250
Regenerative
Specific impulse (s)
¾ Sea level
285
¾ Vacuum
316
Thrust (kN)
¾ Sea level
10 500
¾ Vacuum
Burning time (s)
130
up to 300
Nozzle expansion
ratio
26
81.3
Restart capability
No
No
5.5
No
Yes (up to 8)
Guidance and control
5.5.1 Guidance
Self - contained inertial control system using a precision three-axis gyro-stabilizer and a three-channel voting
on-board digital computer. The digital control computer resides on stage 3. The stage 4 carries an
autonomous digital control system based on three-axis strap down gyros; it also incorporates GPS
navigation systems.
5.5.2 Control
STAGE
Pitch, yaw, roll
1
2
3
4
By gimballing six
nozzles
By gimballing four
nozzles
By 4 verniers
engines (31 kN
thrust)
By 4 thrusters
(396 N thrust) and
12 attitude Control
thrusters
(13.3 N thrust)
Precision
December 2004
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PROTON M
6. DATA SOURCE REFERENCES
1
-
PROTON M/BREEZE M - ILS Tech. Summary Presentation - 1997
2
-
Jane's Space Directory 1998 – 1999, p 249
3
-
PROTON M Launch Vehicle Fact Sheet - ILS - http://www.ilslaunch.com
4
-
PROTON LAUNCH VEHICLE MISSION PLANNER'S GUIDE - ILS - Revision 4 - March 1999
December 2004
Page 18