(National Space Development Agency)

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N-1: (1975) Delta-based orbital launcher built under US license.
N-2: (1981) Uprated vehicle based on N-1.
H-1: (1986) First stage based on N-1/N-2 with Japanese upper stages.
H-2: (1994) New all-Japanese launcher with cryogenic stage 1 & 2 and solid strap-ons.
J-1: (1996) NASDA's first all-solid orbital launcher.


First launch: 3-Feb-1994
Number launched 3 thru May-1996
Launch sites: Yoshinobu Launch Complex, Tanegashima
Principal uses: delivery of 4,000 kg-class payloads into GTO
Vehicle success rate: 100%
LEO (250 km, 30o): 10,000 kg
GTO: 4,000 kg
GEO: 2,000 kg
Sun-synchronous (800 km): 4,300 kg
Escape: 2,800 kg Moon, 2,000 kg Venus, 1,700-2,200 kg Mars (for 3,650 m/s transfer injection), 700 kg Mercury, 400 kg Jupiter
Availability: fishing/range safety requirements currently constrain launches to 45-day periods in Jan/Feb + Aug/Sep each year, plan to have 4/yr capability starting 2000
Cost: Y19 billion; improved H-2 goal is Y14 billion
Number of stages: 2 + 2 strap-ons
Overall length: 50 m with 4S fairing; 51.1 m with 5/4D fairing
Principal diameter: 4.0 m
Launch mass: 260,000 kg + payload for 4S fairing; 278,900 kg + payload for 5/4D fairing + SSBs
Guidance: strapped-down inertial mounted at top of stage 2. NEC inertial guidance computer, 14.2 kg, 45 W consumption, mission reliabilty >0.999, 32 kword main memory capacity (16-bit words), qualified to 12.8 g rms random vibration. Three Japan Aviation Electronics laser gyros + accelerometers

Length: 23.36 m (19.26 m motor)
Diameter: 1.81 m
Mass at ignition: each 70,400 kg
Propellant: 14% HTPB/68% AP/18% Al
Propellant mass: each 59,150 kg
Thrust: each 1,560 kN SL average
Burn time: 94 s
Steering: TVC by ±5o deflection of flexible nozzle joint provides vehicle 3-axis control

The twin Solid Sub-Boosters are carried by stage 1 only when required by mission performance. One flight thru May-1996.
Length: 9.05 m
Diameter: 1.125 m
Mass at ignition: each 10,500 kg
Propellant: 14% HTPB/68% AP/18% Al
Propellant mass: each 8,400 kg
Thrust: each 34.5 kN SL average (SI 267 s vac)
Burn time: 66 s
Steering: none, fixed nozzle canted 10o

Engine: single Mitsubishi LE-7, single start, gimbaled for pitch/yaw control
Length: 28.0 m
Diameter: 4.0 m
Dry mass: 11,900 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 86,200 kg
Thrust: 843.4 kN SL; 1,078 kN vac
Burn time: 346 s
Attitude control: two 1,500 N base auxiliary engines bleed GH2 from LE-7 for roll control following SRB separation.

Engine: single Mitsubishi LE-5A, dual start, gimbaled 3.5o square for pitch/yaw control
Length: 10.6 m
Diameter: 4.0 m
Dry mass: 3,000 kg
Oxidizer: liquid oxygen
Fuel: liquid hydrogen
Propellant mass: 16,700 kg
Thrust: 121.5 kN vac, but 5% idle mode provides precise orbit control + de-orbit capability
Burn time: 609 s max; two burns (403 s to establish LEO parking orbit + 197 s for 28.5o GTO) for typical GEO missions
Attitude control: two IHI hydrazine thruster modules (each 4 x 50 N + 2 x 18 N) provide roll control during burn and 3-axis during coast

The standard 4.1 m diameter, 12.0 m high 1,400 kg aluminum honeycomb core/skin 4S fairing provides a 3.7 m diameter envelope for single payloads. The 4/4D is stretched to 15.0 m for dual main payloads, with an 11 m aluminum upper section and 4 m CEFP lower portion. A 5.1 x 12.0 m version (5S) offers accommodation for Shuttle-class (4.6 m diameter) payloads. The 5/4D is a 14.1 m long dual fairing, with 5.1 m diameter 9.6 m long aluminum upper portion and 4.1 m diameter 4.5 m long CFRP lower. Separation of the clamshell halves is effected by a mild detonating fuse and springs when aerodynamic heating has fallen to 1,135 W/m2 during stage 1 burn.
Acceleration load: peak 4 g longitudinal at stage 1 shutdown, design requirement for 5 g
Vibration: sinusoidal 2 g longitudinal, 1 g lateral
Acoustic load: 141 db peak during liftoff and transonic flight
Thermal environment: 500 W/m2 radiation from inner wall
Pad environment control: 15-25oC air conditioning


The Y65 billion development of the Improved H-2 began in 1995, principally to reduce unit cost from the current Y19 billion to Y14 billion, increase reliability to >0.96 and add a GEO direct injection capability. GTO capacity will be 4,100 kg. First launch will be in 1999. Stage 1 will be little changed, but the LE-7A is being simplified to cut costs. The engine will be throttleable and will provide the LE-7's original thrust specification of 1,180 kN vac when required. The SRBs will change from 4 segments to three. Stage 2 will be extensively reworked, principally separating the main tanks and using the improved LE-5B engine. The current common LOX/LH2 bulkhead is expensive and thermal leakage limits coast durations. Improved H-2 will use tanks separated by a truss structure, reducing cost and allowing a 5 h coast so that the LE-5B can inject payloads directly into GEO. Propellant mass 16,700 kg, thrust increased to 137.3 kN. Hydraulic TVC is replaced by electromechanical actuators.


First launch: 11-Feb-1996 (2-stage suborbital)
Number launched 1 to end-1996
Launch sites: Tanegashima
Principal uses: delivery of intermediate payloads to LEO
Vehicle success rate: 100%
Performance: 870 kg into 250 km, 30o circular
Availability: potentially two annually. Fishing/range safety requirements constrain launches to Jan/Feb + Aug/Sep each year
Cost: about Y5 billion, excluding development cost (not planned commercially)
Number of stages: 3
Overall length: 33.11 m
Principal diameter: 1.81 m
Launch mass: 87,700 kg + payload
Guidance: inertial/radio, largely derived from M-3SII. The B2CNE stage 2 control electronics unit is mounted on stage 2's forward end, and includes the RIG Rate Integrating Gyro on the SFAP Spin Free Analytic Platform for pitch/yaw/roll data. Ground commands are received by B2CNE should deviations from the programmed path be detected. B2CNE routes radio commands and attitude error signals to stage 1's B1CNE, which also uses a stage 1 rate gyro package

Length: 20.99 m
Diameter: 1.81 m
Mass at ignition: 70,900 kg
Propellant: 14% HTPB/68% AP/18% Al
Propellant mass: 59,150 kg
Thrust: 1,560 kN SL average
Burn time: 89 s
Steering: TVC by ±5o deflection of flexible nozzle joint provides vehicle pitch/yaw control. Two aft External Vernier Engine pods ignite 5 s before launch to provide roll control during burn, then 3-axis during coast. EVE 4.18 m long, 56 cm diameter, carrying 110 kg NTO/hydrazine. 3.43-2.45 kN vac blowdown, gimbaled ±45o pitch/roll, ±11o yaw

Designation: M-23
Length: 6.3 m
Diameter: 1.41 m
Mass at ignition: 12,700 kg (excluding fairing)
Propellant: HTPB solid
Propellant mass: 10,400 kg
Thrust: 524 kN vac average
Burn time: 55 s
Attitude control: pitch/yaw control during burn maintained by LITVC, injecting 40 l of NaClO4 solution into nozzle at 8 points. Roll control during burn and 3-axis control in preparation for stage 3 spin-up/separation provided by four clusters of 4 x 150 N Side Jet thrusters at stage 2's base drawing on 40 l hydrazine

Designation: M-3B
Length: 2.7 m
Diameter: 1.495 m
Mass at ignition: 3,590 kg
Propellant: HTPB solid
Propellant mass: 3,280 kg
Thrust: 132 kN vac average
Burn time: 71 s
Attitude control: upper stages spin induced by stage 2 attitude control system prior to separation

The 500 kg glass fiber honeycomb 2 cm thick sandwich clamshell fairing, as used by the M-3SII, is pyrotechnically separated after stage 2 burnout. The 6.86 m long, 1.65 m diameter structure, which also covers stage 3, provides a payload envelope of about 3.0 m long, 1.4 m diameter. Peak vehicle acceleration is 7 g longitudinal (2 g at launch).

Application: H-2 strap-ons boosters
First flown: 3-Feb-1994
Length: 23.36 m total booster, 19.26 m for motor without nose section
Diameter: 1.81 m
Mass at ignition: 70,400 kg
    type: 14HTPB-68AP-18Al
    mass fraction: 0.840 for 59.15 t loading
Propellant mass: 59,150 kg
Burn time: 94 s
Thrust: 1,560 kN SL average
Specific impulse: 273 s vac
Total impulse: about 158 MNs
Expansion ratio: 10:1
Pressure: 46 atm average

Application: H-2 stage 1
First flown: 3-feb-1994
Number flown: 4
Dry mass: 1,714 kg
Length: 3.40 m
Maximum diameter: 1.80 m
Mounting: gimbaled to provide ±7.5o pitch/yaw control, hydraulic pump powered by auxilary turbine mounted on engine
Engine cycle: closed, staged combustion
Oxidizer: liquid oxygen, delivered at 25.0 kg/s in pre-burner and 186.6 kg/s in main chamber
Fuel: liquid hydrogen, delivered at 35.66 kg/s in pre-burner (resultant fuel-rich gas then burned in main chamber with added oxygen)
Mixture ratio: 0.7 in pre-burner; 6.0 main chamber; 5.93 overall
Oxidizer turbopump: 18,285 rpm single-stage centrifugal pump with pre-burner pump; 178.6 atm/253.6 atm main/pre-burner discharge pressures
Fuel turbopump: 41,596 rpm two-stage centrifugal pump; 264 atm discharge pressure
Thrust: 843 kN SL, 1,078 kN vac
Specific impulse: 445.6 s vac
Expansion ratio: 54:1
Combustion chamber pressure: 130 atm main chamber; 206 atm pre-burner
Burn time: 348 s single start

Application: H-2 stage 2
First flown: 3-Feb-1994
Number flown: 4
Dry mass: 244 kg
Length: 2.67 m
Maximum diameter: 1.625 m
Engine cycle: nozzle hydrogen bleed
Oxidizer: liquid oxygen, delivered at 22.83 kg/s total
Fuel: liquid hydrogen, delivered at 5.86 kg/s total
Mixture ratio: 5.0 (20 for 5% idle)
Oxidizer turbopump: 17,363 rpm, 57.8 atm discharge pressure
Fuel turbopump: 50,534 rpm, 63.5 atm discharge pressure
Thrust: 121.5 kN vac; 5% idle mode provides precise orbit control and stage de-orbit
Specific impulse: 452.9 s (>200 s for 5% idle)
Expansion ratio: 130:1
Combustion chamber pressure: 39.3 atm
Burn time: 400 + 210 s dual burn

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