ANGARA 1. IDENTIFICATION. 1.1 Name. 1.2 Classification. Family : ANGARA Series : ANGARA Version : ANGARA 1.1

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1 1. IDENTIFICATION 1.1 Name 1.2 Classification Family : Series : Version : 1.1 Category : SPACE LAUNCH VEHICLE Class : Small Launch Vehicle (SLV) Type : Expendable Launch Vehicle (ELV) 1.3 Manufacturer : Khrunichev SRPSC (State Research and Production Space Center) Novoza Vodskaya ulitsa, 18 MOSCOW Russian Federation (095) Development manager : Khrunichev SRPSC 1.5 Vehicle operator : International Launch Services (ILS) JV Lockheed Martin - Khrunichev - Energia International (LKEI) 1660 International Drive, Suite 800 McLean, VIRGINIA USA Telephone : Fax : Vehicle operator : International Launch Services (ILS) 1.7 Launch cost : 25 M$ (estimation) 2. STATUS 2.1 Vehicle status : Under development 2.2 Development period : First launch : 2005 December 2003 Page 1

2 3. PAYLOAD CAPABILITY AND CONSTRAINTS 3.1 Payload capability Low Earth Orbits The Table 1 below provides the performance capability of the 1.1 launch vehicle for low circular orbits as a function of various orbital inclinations. ORBIT TYPE LEO (kg) Altitude (km) Inclination: TABLE 1: 1.1 LEO PERFORMANCE The dependance of the injected payload mass on target orbit parameters (inclination, altitude) is shown in the Figure 1: FIGURE 1: 1.1 LEO PERFORMANCES FROM PLESETSK December 2003 Page 2

3 3.1.2 Geosynchronous and Interplanetary Orbits No capability Injection accuracy For the 1.1 injecting a payload into an orbit up to 350 km high, the injection accuracy is shown below: ERROR TYPE Injection accuracy 3.2 Spacecraft orientation and separation Orbital altitude ± 2% Inclination ± 0.05 During orbital flight and during coast phases, the 2 nd stage booster can perform programmed turns relative to any of the body axes of the upper stage. The angular velocities of turns relative to any axis do not exceed 1-2 /s. While the main propulsion engine is in operation, control of the booster attitude is determined by the pitch, yaw and roll programs selected for each specific flight program. Any orientation of the booster can be made prior to separation of the S/C. At the time of S/C separation, the booster may be either in stabilization mode or, if necessary, in spinning mode. In stabilization mode, the angular velocities relative to the body axes which can be obtained are: ω X ± 1 /s and ω Y (Z) ± 0.5 /s. The error in the spatial orientation of the booster axes relative to the base inertial coordinate system does not exceed ± 0.5. In spinning mode, an angular velocity relative to the longitudinal axis of the booster of up to 12 /s can be provided. The possibility of raising the angular velocity to 30 /s is being analyzed. 3.3 Payload interfaces Payload compartments and adaptors Payload fairing description: the 1.1 uses the same PayLoad Fairing (PLF) used on the ROCKOT. The mass of the PLF is 710 kg. The structure of the PLF has a composite form. A view of the fairing is presented in Figure 2: FIGURE 2: GENERAL VIEW OF 1.1 PAYLOAD FAIRING December 2003 Page 3

4 Payload adaptors and separation system: the S/C adaptor system is used to mount the S/C on the mm diameter of the interface with the launcher when the S/C design does not allow it, where the openings are located on the mating ring of the L/V. The shape and height of the adaptor system are selected in relation with the structural features. The Figure 3 shows different views of the payload adaptor system. FIGURE 3: PAYLOAD ADAPTOR December 2003 Page 4

5 Depending on the requirements of the customer, a separation system with a standard size of or mm or with some other size specified by the customer may be used. Spring pushers are mounted on the top end ring to provide the initial impulse upon separation of the S/C from the L/V. The number of spring pushers and their forces are determined by the S/C separation requirements. A total of up to 12 spring pushers may be installed on the adaptor system. Two brackets with umbilical electrical connectors that provide electrical connection between the S/C and the L/V are installed on the adaptor system structure. The type of umbilical electrical connectors, and the coordinates at which they are installed, are determined in accordance with the Customer. Payload constraints: during the ascent phase of flight, the S/C is in radiative heat transfer with the inside surface of the PLF. The allowable temperature level of the PLF structure is maintained by application of a thermal protective material. A thermal insulation material lined with a film with low radiant emissivity (< 0.1) is mounted on the inside surface of the PLF. The value of the maximum radiant heat flux from the inside surface of the PLF to the S/C will not exceed 250 W/m 2 from the launch time until the PLF separation. Based on a separate requirement by the S/C manufacturer, the level of radiant heat flux to the S/C can be lowered to W/m 2. In absolute value, the maximum decrease rate of the pressure under the PLF does not exceed 5 kpa/s. 3.4 Environments Mechanical environment The maximum (static and dynamic) loads on the S/C at launch and during flight of the 1.1 are presented in Table 2. Transverse loads may act in any direction perpendicular to the longitudinal axis of the L/V. LOADING CASE SAFETY FACTOR LONGITUDINAL LOADS (g) TRANSVERSE LOADS (g) STATIC DYNAMIC QUASI-STATIC Launch ± 0.7 ± 1 Flight at Q max ± 0.4 Flight at P max ± 0.9 Flight at n xxmax ± 0.4 ± 0.6 Separation of 1 st and 2 nd stages ± 0.46 ± nd stage flight ± 0.2 TABLE 2: MAXIMUM QUASI-STATIC LOADS AT LAUNCH AND DURING FLIGHT Acoustic vibrations Low frequency vibration - Vibration loads on S/C are presented in the form of wide-band random vibrations and sinusoidal vibrations. December 2003 Page 5

6 - Vibration loads on S/C in the form of sinusoidal vibrations are presented in Tables 3 and 4: FREQUENCY RANGE (Hz) VIBRATION ACCELERATION (m/s 2 (g)) ACTION TIME (s) (1-3) (3-10) (10) 240 TABLE 3: HARMONIC VIBRATION LOADS DURING 1 st STAGE OPERATION FREQUENCY RANGE (Hz) VIBRATION ACCELERATION (m/s 2 (g)) ACTION TIME (s) ( ) (1.44-6) (6) 706 TABLE 4: HARMONIC VIBRATION LOADS DURING 2 nd STAGE OPERATION - Vibration loads on S/C in the form of wide-band random vibrations are presented in Table 5: EQUIPMENT INSTALLATION POINT FREQUENCY RANGE (Hz) SPECTRAL DENSITY OF VIBRATION ACCELERATION (m 2 -s -4 /Hz (g 2 /Hz)) ACTION TIME (s) Interface between L/V and S/C during 1 st stage operation (0.068) 7.7 (0.08) 5.54 (0.058) 5.13 (0.053) 6.47 (0.067) 4.28 (0.044) 2.14 (0.022) 240 Interface between L/V and S/C during 2 nd stage operation 2.13 (0.022) 1.76 (0.018) 1.92 (0.02) 5.55 (0.058) 2.57 (0.027) 1.53 (0.016) (0.008) 706 TABLE 5: RANDOM VIBRATIONS DURING 1 st AND 2 nd STAGES OPERATIONS The vibration and shock loads, in the direction of all three axes at stage separation, are presented as the values of the shock spectrum in Table 6: FREQUENCY SUB-RANGE (Hz) VALUE OF SHOCK SPECTRUM (m-s -2 /Hz (g 2 /Hz)) (25-50) (50-150) ( ) ( ) ( ) (5 000) TABLE 6: VIBRATION AND SHOCK LOADS AT STAGE SEPARATION December 2003 Page 6

7 Acoustic pressure The acoustic loads under the payload fairing do not exceed the values presented in the following Figure 4 and Table 7. FIGURE 4: ACOUSTIC LOADS UNDER FAIRING CENTER FREQUENCY OF 1/3 - OCTAVE FREQENCY BAND (Hz) ACOUSTIC PRESSURE (db) Total acoustic pressure level, db Duration 60 s TABLE 7: ACOUSTIC LOADS UNDER FAIRING December 2003 Page 7

8 3.4.3 Shock The vibration and shock loads upon separation of the S/C (preliminary values) in the directions of the three axes are presented in the form of the shock spectrum values in Table 8: FREQUENCY SUB-RANGE (Hz) VALUE OF SHOCK SPECTRUM (m-s 2 (g)) ,200 17,200-49,000 49,000 (25-50) ( ) ( ) ( ) ( ) (5 000) Thermal environment TABLE 8: SHOCK LOADS AT S/C SEPARATION Ground thermal loads on the S/C arise during transportation of the S/C and during launch processing of the S/C at the technical and launch areas. Information on the environmental parameters around the S/C during various phases of ground processing of the S/C for launch, and on the means used to maintain them, are presented in Table 9. PROCESSING PHASE Air transportation to Cosmodrome Motor vehicle transportation to technical complex Transportation of S/C as part of L/V to launch complex S/C processing and filling Integration with upper stage and L/V Erection at launch complex Processing at launch complex TEMPERATURE ( C) HUMIDITY (%) Meets the conditions for transportation in an unsealed cabin MEANS USED TO MAINTAIN Container Container Thermal control unit Technical systems Technical systems Launch abort TABLE 9: ENVIRONMENT PARAMETERS AROUND S/C Air thermal control system (thermal control through transporter/erector). Flow up to kg/h. Air thermal control system (air supply through cable and filling tower). Flow up to kg/h. Air thermal control system (air supply through cable and filling tower until transporter/erector is connected; flow is up to kg/h) until mobile thermal control unit is brought up. December 2003 Page 8

9 Conditions in the technical complex premises Temperature Between 10 C (± 2 ) and 30 C (± 2 ) Relative humidity From 30% to 60% Cleanliness Class at air inlet payload compartment Conditions during the flight under the fairing During the ascent phase of flight, the S/C is in radiative heat transfer with the inside surface of the PLF. The allowable temperature level of the PLF structure is maintained by application of a thermal protective material. A thermal insulation material lined with a film with low radiant emissivity (ε < 0.1) is mounted on the inside surface of the PLF. The value of the maximum radiant heat flux from the inside surface of the PLF to the S/C will not exceed 250 W/m 2 from the time of launch until the PLF separates. Based on a separate requirement by the S/C manufacturer, the level of radiant heat flux to the S/C can be lowered to W/m Operation constraints Ground constraints : N/A Launch rate capability : - Procurement lead time : - 4. LAUNCH INFORMATION 4.1 Launch site Plesetsk The ground complex is located at the Plesetsk Cosmodrome in Arkhangelsk Oblast, about 800 km northnortheast of MOSCOW. The geographic coordinates of the Cosmodrome are 62.7 North, 40.3 East. The climatic conditions of the Cosmodrome are characterized by the following data: mean air temperature : - 28 C in winter, + 23 C in summer, maximum precipitation rate : 0.04 mm/mn, mean annual precipitation : 398 mm. The is designed for reliable operation in this range of climatic conditions. December 2003 Page 9

10 Launch complex FIGURE 5: PLESETSK LAUNCH SITE - SCHEME OF MAIN FACILITIES The universal launch complex is the aggregate of the launch facilities, engineering systems, and handling and support equipment and is designed to receive the, perform pre-launch processing, and carry out launch. The delivery of the L/V to the launch pad, transfer of the L/V to the vertical position, and mounting of the L/V on the launch pad are carried out by the transporter/erector. In the erected position, the 1 st stage of the L/V is connected to ground handling equipment through the automatic umbilical mating units, which are housed in the launch pad structure, and connected to the receptacles on the bottom of the L/V. The L/V processing at the ground complex is based on the principle of parallel processing of its components (L/V stages, upper stage and S/C). During processing phases, preliminary operations, tests, checks, and filling of L/V tanks with propellants and compressed gases are carried out at the appropriate technical complex facility. The final operations and testing of the L/V and filling of L/V propellant tanks are carried out at the universal launch complex. 4.2 Sequence of flight events Two payload injection schemes may be used for launches from the Plesetsk Cosmodrome, depending on the altitude of the targeted orbit. Direct Injection Scheme: after separation of the 1 st stage booster, the L/V 2 nd stage booster, with a single burn, inserts the payload into the specified orbit. This scheme is implemented for circular orbits with altitudes up to 300 km. Scheme Using the Coast Phase of the L/V 2 nd stage: after the 1 st stage booster separation, the 1 st burn of the main propulsion engine of the 2 nd stage injects the payload into an elliptical transfer orbit with an apogee altitude equal to the altitude of the specified circular orbit. Until apogee of the transfer orbit is reached, the 2 nd stage booster is in a coast phase. At orbital apogee, a 2 nd burn of the 2 nd stage main propulsion engine transfers the payload into the targeted orbit. When the payload is inserted into orbits with an altitude of from 300 to km, the coast phase time of the 2 nd stage changes from 45 to 51 min. December 2003 Page 10

11 When optimizing a trajectory design, the program for the pitch angle of the L/V 1 st and 2 nd stages is selected so that the jettisoned hardware (the 1 st stage booster and PLF) will fall into specified regions. The time of jettisoning of the PLF is determined so that the fairing pieces fall into the same region as the 1 st stage booster. The maximum value of the free molecular heat flux acting on the S/C after PLF jettison does not exceed W/m 2. Figures 6 and 7 show a typical flight scheme, possible inclinations of injection orbits and impact regions of jettisoned hardware. FIGURE 6: SEQUENCE OF FLIGHT EVENTS FIGURE 7: 1.1 POSSIBLE ORBIT INCLINATIONS AND JETTISONED HARDWARE IMPACT ZONES December 2003 Page 11

12 The L/V ascent characteristics for a payload directly injected into a circular orbit, 200 km altitude with an inclination of 63, is presented in Figure 8: 4.3 Launch record data FIGURE 8: 1.1 ASCENT CHARACTERISTICS LAUNCH DATE NUMBER OF SATELLITES ORBIT RESULT REMARK None Previsional reliability : - Failure : - Success ratio : Planned launches 1 st launch is foreseen in 2005 December 2003 Page 12

13 5. DESCRIPTION 5.1 Launch vehicle FIGURE 9: 1.1 LAUNCH VEHICLE 5.2 Overall vehicle Overall length Maximum diameter Lift-off mass : m : 2.9 m : 149 t December 2003 Page 13

14 5.3 General characteristics of the stages STAGE 1 2 Designation CRM 1 (Common Rocket Module) or URM 1 (Universal Rocket Module) CRM 2 or URM 2 Manufacturer Khrunichev Khrunichev Length (m) Diameter (m) Dry mass (t) - - Propellant: Liquid Liquid Type Storable Storable Fuel Kerosene Kerosene Oxidizer LO 2 LO 2 Propellant mass (kg) Fuel - - Oxidizer - - Tank pressurization - - Tank presure (bar) - - Lift-off mass (t) December 2003 Page 14

15 5.4 Propulsion STAGE 1 2 Designation CRM 1 or URM 1 CRM 2 or URM 2 Engine RD-191M RD-0124A Manufacturer NPO Energomash Number of chambers 1 4 Engine mass (kg) Length (m) Diameter (m) 3.8 nozzle exit section: at thermal protective screen: 2.4 Feed system type - - Mixture ratio - - Chamber pressure (bar) Cooling - - Propellant Feed System Closed-cycle turbopump Closed-cycle turbopump Verniers : single shared turbopump Specific impulse (s) Sea level Vacuum Thrust (kn) Sea level Vacuum Burning time (s) Nozzle expansion ratio - - Throtting Capability - - Restart capability No Up to 8 FIGURE 10: RD-191 AND RD-0124A ENGINES December 2003 Page 15

16 5.5 Guidance and Control The control system of the family is based on the well-refined design solutions used in the control system of the Proton M and the Breeze M. The control system equipment and units in the CRM are common to the entire L/V family. In the 2 nd stage of the L/V, the control system is common with the control system of the Proton M. The on-board control system is self-contained and inertial, and is based on the on-board digital computer system. The structure of the control system uses the modular method for communications between devices. This method allows both system growth and updating of individual control system while preserving the minimum number of connections between them. 5.6 Launch Vehicle growth The underlying concept for the family is a modular construction, using a booster core named Universal Rocket Module (URM) or Common Rocket Module (CRM). The CRM forms the 1 st stage booster of the light class 1.1 and 1.2, which differ in their 2 nd stage boosters. The medium and heavy class launch vehicles, respectively, consist of three or five CRMs. An illustration of the design concept for the family is presented in the Figure A3 A5 A5/KVRB FIGURE 11: COMPOSITION AND DESIGNATION OF FAMILY December 2003 Page 16

17 The Tables 10 to 12 present the main characteristics of the different versions: 1.2 A3 A5 A5/KVRB Launch weight (t) Maximum length (m) Maximum diameter (m) Control system Inertial TABLE 10: FAMILY CHARACTERISTICS Mass of propellant (t) Propellant components Propulsion system thrust (metric tons-force) sea level vacuum LO2 + RG-1 TABLE 11: 1 st STAGE BOOSTER Mass propellant and gases (t) Propellant components LO2 + RG-1 LO2 + RG-1 LO2 + RG-1 LO2 + H2 Propulsion system thrust (in vacuum) (metric tonsforce) TABLE 12: 2 nd STAGE BOOSTER To satisfy as much as possible the Customer s requirements, the family is compatible with three upper stage options, whose characteristics are presented in the Table 13. BREEZE M CORE BREEZE M WITH ADDITIONAL PROPELLANT TANK KVRB (LO 2 - LH2) UPPER STAGE Lift-off mass (t) Propellant mass (t) Thrust (t) Dimensions (d l) (m) TABLE 13: UPPER STAGE GENERAL CHARACTERISTICS December 2003 Page 17

18 6. DATA SOURCE REFERENCES 1 - Launch System Mission Planner s Guide - December Revision 0 (ILS) 2 - International Reference guide to Space Launch Systems - 3 rd Edition - S.Isakowitz - AIAA 3 - Jane's Space Directory 4 - IAA -99-IAA Novosti Kosmonavtiki, n 3 (194) Page "" Launch vehicle family concept, development status and operational plans A. Medvedev, A. Kuzin, E. Motorny, Khrunichev Space Center, RUSSIA B. Katorgin, NPO Energomash, RUSSIA December 2003 Page 18

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