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    Displaying results 1 - 4 of 14
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  • Proton M/BREEZE M

    Proton M/BREEZE M

    International Launch Services

    Russia

    Price negotiable

    • Performance Summary

      Adapter mass must be subtracted to determine separated spacecraft mass. Performance reflects commercial configuration and flight profile. Maximum performance for Russian government missions is higher. See Performance section for more details.

      180 km (97.2 nmi), 51.5 deg 23,700 kg (52,250 lbm)
      185 km (100 nmi), 90 deg ?
      Space Station Orbit: 186×222 km (100×120 nmi), 51.6 deg 23,700kg (52,250 lbm). The three-stage Proton delivers spacecraft to low orbit, then the spacecraft must use onboard propulsion to raise the orbit to space station altitude.
      Sun-Synchronous Orbit: 800 km (432 nmi), 98.6 deg ?
      GTO: 1500 m/sec Delta-V to 650 6300 kg (13,889 lbm)
      Geostationary Orbit 3300 kg (7275 lbm)

      Flight Record (through 31 July 2021)
      Total Orbital Flights 112
      Launch Vehicle Successes 101
      Launch Vehicle Partial Failures 2
      Launch Vehicle Failures 9

      Proton is launched from the Baikonur Cosmodrome in Kazakhstan. Because Baikonur is a landlocked launch site, specific drop zones are reserved for the impact of the first stage, second stage, and payload fairing. These drop zones are typically about 310 km (190 mi) downrange for the first stage, and 1985 km (1230 mi) downrange for the second stage and payload fairing. Proton is therefore constrained to fly along one of three launch azimuths that are aligned with these zones. The basic three-stage Proton does not have an upper-stage restart capability, and therefore cannot reach orbits with perigees higher than approximately 200 km (108 nmi). To reach higher orbits or other inclinations, Proton delivers the payload plus an upper stage to a standard 180–200 km (97–108 nmi) circular support orbit. The upper stage or spacecraft propulsion system then performs orbit raising and/or plane change maneuvers. (While the azimuth restrictions limit performance, this approach does have the advantage of standardizing the trajectory and flight software of the lower three stages.)

      Historically the Block DM upper stage has been used for most Proton K flights. Performance of the three-stage Proton configurations to typical support orbits is provided subsequently. Performance varies depending on the fairing separation time. Two separation times are available, each of which meets impact point restrictions. Payload fairing separation at 185 s results in higher performance than the 350 s separation time, but also exposes the payload to heating rates much higher than are standard for Western spacecraft. An intermediate time is used by Proton M to enable full performance for Western satellite missions. Performance for Russian government missions, which may use smaller payload fairings and the early fairing jettison time, can be higher than the commercial advertised performance. A comparison is shown below.

        Proton M/Breeze M  
        Commercial Russian Government
      180 km (97.2 nmi), 51.5 deg 22,000 kg (48,400 lbm) 23,700 kg (52,250 lbm)
      Geostationary Orbit (GSO) 3200 kg (7040 lbm) 3300 kg (7275 lbm)

      The Block DM is a LOX/kerosene-fueled stage with a capability for up to seven engine restarts (five have been demonstrated) and an on-orbit lifetime of 24 h, or longer with additional modifications. This allows the Block DM to deliver payloads directly to GSO or to coast for several revolutions in the support orbit for proper phasing and then deliver a spacecraft to GSO close to its planned longitude. Russian domestic satellites have generally been delivered directly to GSO, but commercial spacecraft are typically delivered to a GTO with the apogee and perigee optimized for the particular mission. For certain GTO orbits, a three-burn trajectory can be used in which the Block DM stage is delivered to a suborbital trajectory and performs its first burn to reach the parking orbit. This results in higher performance as shown subsequently. Performance is increased for certain domestic Russian missions by using the lighter Block D2 stage, which lacks the avionics compartment of the Block DM. This results in a performance improvement of approximately 800 kg (1750 lbm), but requires that the spacecraft be capable of controlling the stage. This configuration is used primarily for Russian planetary spacecraft.

      The Block DM is also used for many Proton LEO missions (for example, on three Iridium deployment flights), to deliver the payload to the desired inclination and altitude. However, the Block DM has propellant loading constraints that impact its performance for LEO and other low-energy missions. A minimum of 11,500 kg (25,350 lbm) of propellant must be loaded onto the Block DM, even for missions that do not require this much capability from the stage. As a result of carrying this surplus propellant, the available payload mass is limited and is largely independent of altitude for LEO missions. Missions to higher altitudes simply burn more of the propellant load, leaving a smaller surplus.

      The new Breeze M is a modified version of the Breeze KM upper stage developed for the Rockot space launch vehicle. Like the Block DM, the Breeze M is capable of 24 h of on-orbit operations and multiple restarts. However, the Breeze M main engine has a much lower thrust (and therefore a longer burn time) than the Block DM. As a result, the apogee raising maneuver for GTO missions is usually, but not always, split into two burns. The first puts the Breeze M and payload into an intermediate elliptical orbit, and the second raises the apogee to the planned altitude.

      Performance values shown are for commercial payloads. The standard commercial payload fairing is assumed for Block DM missions. The short version of the Breeze M fairing is assumed. The larger Breeze M payload fairing reduces payload capacity by roughly 100 kg (220 lbm) for GTO missions. The later of two available fairing separation times is used to keep free molecular heating rates at acceptable levels for commercial payloads. Domestic Russian payloads typically use smaller fairings, and in some cases separate the payload fairing earlier, resulting in higher performance than shown here.

      The performance shown is the payload systems mass, which includes both the spacecraft mass and the mass of the payload adapter and any additional payload support hardware such as cable harnesses. Typical mass for Block DM adapters is 110–155 kg (243–342 lbm). Typical mass for Breeze M adapters is 110 kg (240 lbm). Sufficient flight performance reserve is included to provide 3 sigma confidence of performance levels.

      Three-Stage Proton Performance
      Orbit Proton M/Briz M
      180 km (97.2 nmi), 51.5 deg 22,000 kg (48,400 lbm)
      190 km (103 nmi), 64.8 deg 20,610 kg (45,435 lbm)
      170 km (92 nmi), 72.7 deg 19,975 kg (44,035 lbm)

      Performance

      Proton-Performance-Graphic

    • Payload Compartment

      Max Payload Diameter:
      3864 mm (152.1 in.)
      Max Cylinder Length:
      BR-1160: 4112 mm (161.9 in.); BR-13305: 4116 mm (162.0 in.); BR-15255: 6064 mm (238.8 in.)
      Maximim Cone Length:
      3800 mm (149.6 in.)
    • Dimensions

      Length:
        Stage 1 Stage 2 Stage 3 Stage 4
      Block DM
      Stage 4
      Breeze-M
      Dimensions          
         Length 21.2 m (69.5 ft) with interstage: 17.1 m (55.9 ft) 6.9 m (22.6 ft) 6.3 m (20.7 ft)
      without payload adapter
      2.8 m (9.2 ft)
      Diameter:
        Stage 1 Stage 2 Stage 3 Stage 4
      Block DM
      Stage 4
      Breeze-M
      Dimensions          
         Diameter Core: 4.1 m (13.4 ft)
      External tanks: 1.7 m (5.6 ft)
      Total: 7.4 m (24.3 ft)
      4.1 m (13.4 ft) 4.1 m (13.5 ft) 3.7 m (12.1 ft) 4 m (13 ft)
  • Proton K/BLOCK DM

    Proton K/BLOCK DM

    International Launch Services

    Russia

    Retired.

    • Performance Summary

      Adapter mass must be subtracted to determine separated spacecraft mass. Performance reflects commercial configuration and flight profile. Maximum performance for Russian government missions is higher. See Performance section for more details.

      200 km (108 nmi), 51.6 deg 19,760 kg (43,560 lbm)
      185 km (100 nmi), 90 deg 3620 kg (7980 lbm)
      Space Station Orbit: 186×222 km (100×120 nmi), 51.6 deg 19,760 kg (43,560 lbm). The three-stage Proton delivers spacecraft to low orbit, then the spacecraft must use onboard propulsion to raise the orbit to space station altitude.
      Sun-Synchronous Orbit: 800 km (432 nmi), 98.6 deg 3620 kg (7980 lbm)
      GTO: 1500 m/sec Delta-V to 650 4350 kg (9590 lbm)—two upper-stage burns
      4930 kg (10,846 lbm)—three upper-stage burns
      Geostationary Orbit 1880 kg (4145 lbm)
    • Payload Compartment

      Max Payload Diameter:
      Max Cylinder Length:
      Maximim Cone Length:
    • Dimensions

      Length:
      Diameter:
  • Pegasus

    Pegasus

    Orbital Sciences Corporation

    United States

    $15–25 million (OSC, 2002)

    • Performance Summary

      The performance for the Pegasus launch vehicle with a 965-mm (38-in.) marmon clamp interface is shown below.

      Available Inclinations: 0–180 deg
      185 km (100 nm), 28.5 deg 443 kg (977 lbm)
      185 km (100 nm), 90 deg 332 kg (732 lbm)
      Space Station Orbit: 407 km (220 nm), 51.6 deg 361 kg (796 lbm)
      Sun-Synchronous Orbit: 800 km (432 nm), 98.6 deg 190 kg (420 lbm)

      Pegasus missions begin when the Orbital carrier aircraft (OCA) releases the Pegasus booster at an altitude of approximately 12 km (39,000 ft) and a speed of Mach 0.8. The added initial velocity and lower drag of an air launch allow Pegasus to be smaller than a booster of equivalent capability launched from the ground. In addition, the air launch capability means Pegasus can be launched from any ocean area with suitable tracking facilities, and therefore it can reach almost any orbit inclination. In a typical mission, the first-stage booster ignites 5 s after release from the aircraft and quickly initiates a pull-up maneuver to gain altitude. The vehicle decreases its pitch angle about 25 s into flight, and the second-stage ignition occurs shortly after the first stage burns out. There is a long coast period before third-stage ignition to allow the vehicle to reach orbital altitude, and then the third stage provides the remaining velocity to achieve orbit. Because solid motors have relatively large uncertainties in performance, Pegasus trajectories are designed with an additional velocity margin in the form of a flight performance reserve. If motor performance is less than nominal, the velocity margin ensures sufficient performance to reach the target orbit. If motor performance meets or exceeds nominal levels, the additional velocity can be eliminated by steering out of plane, or it can be used to maximize the orbit altitude. Orbital offers a small hydrazine-fueled fourth stage called the hydrazine auxiliary propulsion system (HAPS) located inside the avionics section for minimal impact on available payload volume. Missions using the HAPS stage can achieve higher performance at high altitude and also have improved orbit injection accuracy.

      Pegasus-Performance-Graphic

    • Payload Compartment

      Max Payload Diameter:
      1168 mm (46.0 in.)
      Max Cylinder Length:
      1110 mm (43.72 in.)
      Maximim Cone Length:
      1016 mm (40.0 in.)
    • Dimensions

      Length:
      Stage 1
      Orion 50SXL
      Stage 2
      Orion 50XL
      Stage 3
      Orion 38
      Stage 4
      HAPS (Optional)
      10.3 m (33.8 ft) 3.11 m (10.2 ft) 1.34 m (4.4 ft) 0.71 m (2.3 ft)
      Diameter:
      Stage 1
      Orion 50SXL
      Stage 2
      Orion 50XL
      Stage 3
      Orion 38
      Stage 4
      HAPS (Optional)
      1.28 m (4.2 ft) 1.28 m (4.2 ft) 1 m (3.2 ft) 1.0 m (3.3 ft)
  • Minotaur-C

    Minotaur-C

    Northrop Grumman Space Systems

    United States

    $40-50 million (OSC, 2014)

    • Performance Summary

      200 km (108 nmi), 28.5 deg 14578 kg (3214 lbm)
      200 km (108 nmi), 90 deg  
      Space Station Orbit: 407 km (220 nm), 51.6 deg ?
      Sun-Synchronous Orbit: 800 km (432 nm), 98.6 deg 1054 kg (2324 lbm)
      GTO: 185×35,786 km (100×19,323 nm), 28.5 deg  
      GEO No capability

      Flight Record (through 30 June 2021)
      Total Orbital Flights 10
      Launch Vehicle Successes 7
      Launch Vehicle Partial Failures 0
      Launch Vehicle Failures 3

      Minotaur-C-Graphic-1

      Orbital Flights Per Year

      Minotaur-C-Graphic-2

      Minotaur-C-Performance-Graphic3

      Minotaur-C Launches 2004-2021

      Failure Descriptions:      
      P 1998 Feb 10     T2 1998 007 Delivery orbit apogee was 91 km higher than planned. Although considered a partial failure according the definition used in this publication, both Orbital Sciences and the payload customer consider the launch a success.
      F

      ADD: Minotaur-C 2004-2021 failures and partial failures (spreadsheet) here (once available)
      2001 Sep 21     T6 2001 F01 When the second stage ignited at T+83 seconds, a nozzle gimbal actuator drive shaft siezed
      for approximately 5 seconds causing loss of control. The vehicle recovered and continued to fly
      the mission profile, but failed to reach a stable orbit and reentered near Madagascar.
    • Payload Compartment

      Max Payload Diameter:
      Max Cylinder Length:
      Maximim Cone Length:
    • Dimensions

      Length:
      Stage 0 Stage 1 Stage 2 Stage 3
      12.8 m (41.9 ft) 8.6 m (28.3 ft) 3.1 m (10.1 ft) 1.3 m (4.4 ft)
      Diameter:
      Stage 0 Stage 1 Stage 2 Stage 3
      2.35 m (7.7 ft) 1.27 m (4.17 ft) 1.27 m (4.17 ft) 1.0 m (3.17 ft)
    Displaying results 1 - 4 of 14
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