Chang Zheng-2C (Long March-2C)
- Name: Chang Zheng-2C (CZ-2C)
- English Name: Long March-2C (LM-2C)
- Type: Orbital launch vehicle
- Developer: CALT
- First Flight: 1975
- Status: Operational
- Variants: 2C, 2C Batch-02, 2C/SD, 2C+, 2C/SM, 2C/SMA
- Launch Sites: Jiuquan, Taiyuan, Xichang
- Hai Yang 1
- Huan Jing 1
- Shi Jian 11
- Tan Ce 1
- Yaogan Weixing
The Chang Zheng-2 (CZ-2, or “Long March-2”) was the space launcher version of the two-stage liquid-propellant Dong Feng-5 (DF-5, CSS-4) intercontinental ballistic missile. It has also been used as the base rocket for the subsequent CZ-3 and 4 launch vehicle families. Over the years, the CZ-2 itself has been modified with extended propellant tank and the addition of various upper stages for different mission profiles.
In the late 1960s, the 1st Academy (now CALT) of the Seventh Ministry of Machinery Industry was tasked with the development of a two-stage, liquid-propellant launch vehicle to loft the 1,800 kg FSW (Jian Bing 1) recoverable satellite. The 1st Academy, based in Beijing, produced two design proposals, both based on the DF-5 ICBM. The first design would replace the YF-23 engine on the second stage of the DF-5 with a newly-developed YF-25 liquid engine with 25 t thrust, which would increase the rocket’s payload capacity by 500—800 kg. The second proposal retained the DF-5 design but focused on optimising the rocket’s flight profile by extending the burn time of the second stage vernier motor, which would increase the payload capacity by 500 kg.
Space programme planners asked both designs to be developed in parallel, but the delay in the development of the YF-25 engine meant that only the second proposal could meet the scheduled first launch in 1974. The YF-25 engine finally passed 130 sec ground test in 1979, but the idea to adopt it on the launch vehicle was abandoned.
The DF-5 ICBM made its maiden flight from the Jiuquan Satellite Launch Centre in northwest China on 10 September 1971, but the test was only partially successful. The second test on 8 April 1973 ended up with the rocket exploding in the midair 43 seconds after liftoff. The DF-5 development was subsequently suspended under the order of Premier Zhou Enlai. The remaining three test missiles were modified into the CZ-2 launch vehicles for the FSW launch missions.
On 5 November 1974, a CZ-2 rocket carrying an FSW satellite was launched from Jiuquan but exploded in the midair only 20 seconds into flight. Subsequent investigations revealed that the failure was caused by a disconnected cable for the putch rate gyro signal in the launch vehicle’s guidance system.
On 26 November 1975, a second CZ-2 launch was successful, with the FSW satellite placed into a 185 km LEO. This was followed by further two successful launches in December 1976 and January 1978.
In 1979, the Chinese military ordered six more CZ-2 rockets for subsequent FSW missions. As the rockets from the Barch-01 production of the DF-5 were all used up, the new rockets would be built by converting the missiles in the Batch-02 production of the DF-5. The 1st Academy used the opportunity to introduce a number of improvements to the launcher’s design and this improved variant was designated CZ-2C (the 2A and 2B designations had already been allocated to the two three-stage launch vehicles under development at the time, which later became CZ-3 and CZ-4).
In February 1982, the FSW satellite and CZ-2C launch vehicle were officially certified by the COSTIND (Commission of Science, Technology and Industry for National Defence) for batch production, with costs for future launch missions covered by miltiary budgets.
The CZ-2C made its maiden flight on 9 September 1982, successfully placing the FSW-0-04 satellite into orbit and achieving a record high accuracy in orbit insertion. Five more launches were conducted between August 1983 and August 1987, all of which were successful. The 1987 flight also carried two French microgravity experiment packages on the piggyback – the first ever orbital launch service for a foreign customer.
To support the launch of the heavier FSW-1 (Jian Bing 1A) satellite, CALT made further improvements to the CZ-2C design in the mid-1980s, including adding it with the telemetry system of the DF-5. To distinguish it from the previous CZ-2C, this variant was internally designated CZ-2C Batch-02. The rocket made its maiden flight on 9 September 1987, placing the 2,076 kg FSW-1-01 into orbit.
In 1993, the Chinese space industry signed a contract with the U.S. telecommunication firm Motorola to use its CZ-2C rocket to loft 12 satellites for the Iridium global wireless communications satellite network. A new launcher designated CZ-2C/SD was specifically developed to support these missions. Modifications on the CZ-2C/SD included a 3.35 m diameter payload fairing, a stretched second stage for additional propellant, and an improved second stage engine with higher expansion ratio nozzles. The rocket’s overall length was increased from 31.17 m to 40 m. Its liftoff weight was increased from 192 t to 213 t. Its payload capacity to LEO was increased from 2,400 kg to 3,000 kg.
To place two satellites in a single launch mission, the launch vehicle was added with a three-axis stabilised, solid rocket upper stage known as “Smart Dispenser” (SD). Developed by CALT, the upper stage consisted of a 742 kN thrust main motor and four 74.1 kN venire motors, with its own onboard guidance system.
Between 1997 and 1999, the CZ-2C/SD made 7 flights from the Taiyuan Satellite Launch Centre, lofting two mock satellites and 12 operational Iridium satellites.
In the early 2000s, CALT made further improvements to the CZ-2C in order to support the launch of the 4,000 kg mass FSW-4 (Jian Bing 2) recoverable satellite. The overall length of the rocket was stretched by 3 m to accommodate additional propellant. The rocket made three successful flights between 2004 and 2006, placing two FSW-4 satellites and a Shi Jian 8 satellite into orbit.
To support the launch of the Tan Ce scientific research satellites under the “Double Star” Programme, the CZ-2C was added with a spin-stabilised solid motor kick stage designated SM. The first flight took place on 30 December 2003 from the Xichang Satellite Launch Centre, followed by the second flight on 25 July 2004 from the Taiyuan Satellite Launch Centre.
On 6 September 2008, a CZ-2C placed two satellites, Huan Jing 1A and 1B, into their intended orbits. This rocket was fitted with a new 3-axis stabilised solid motor kick stage designated SMA.
On 18 August 2011, a CZ-2C carrying the Shi Jian 11-04 satellite failed to deliver the payload into its intended orbit due to a malfunction of the launcher.
The first stage of the rocket comprises (from front to rear): interstage structure, oxidiser tank, inter-tank section, fuel tank, engine frame, and tail section. The stage has a body diameter of 3.35 m, which is the maximum to allow it to fit through railway tunnels for transport from its factory to the launch sites.
The two propellant tanks form part of the thrust and weight bearing load structure and are connected by an inter-tank section. The propellant tanks and the inter-tank section are all 3.35 m in diameter and constructed from high-strength aluminium-alloy LD10. The front end of the forward oxidiser tank is protected by a fibreglass heat insulation layer to prevent damage due to the high pressure and hot stream of engine exhaust from the second stage engine during stage separation. Oxidiser is pumped to the main engines via a pipe penetrating through the centre of the rear fuel tank.
The four engines are mounted on the engine frame secured to the rear of the fuel tank. The frame transfers the thrust of the engines to the rocket’s thrust and weight bearing load structure.
The tail section is a two-piece shroud 3.5 m in diameter and 2.4 m in length. It protects the main engines from aerodynamic forces and also houses engine components and linkages. On some variants, the tail section also incorporates four fixed stabilising fins. A honeycombed fibreglass heat shield situated between the engine nozzles and the rear structure safeguards engine components and the fuel tank from the flames and heats of engine exhaust during launch. Four pressure relief valves regulate the pressure inside the shroud during flight.
When the launch vehicle is erected on the launch pad, the vehicle is supported by four weight bearing points at the base of the stage. The weight bearing points are secured to the launch pad’s base unit with explosive bolts, which are detonated less than a second before the launch vehicle lifts off.
The first stage is connected to the second stage with an interstage structure comprising an interstage shroud and a grid structure. The interstage shroud, which is 3.35 m diameter and 3.2 m in length, houses the second stage engines to protect them from aerodynamic forces and is connected to the rear of the second stage fuel tank with explosive bolts. The grid structure, which consists of thirty-two 60 mm diameter metal bars, was designed to let exhaust gas from the engines on the second stage to escape.
The rocket uses a ‘hot separation’ method where the two stages are separated by the impingement of the hot exhaust gas jet from the second stage engines. The engines ignite while the two stages are still connected together, thus eliminating the need for jettisoning devices to provide the separation impulse and considerably reducing the duration of unpowered flight during the separation.
The second stage maintains the overall diameter of the first stage at 3.35 m and comprises (from front to rear): instrument compartment, oxidiser tank, inter-tank section, and fuel tank. The propellant tanks are constructed from high-strength aluminium-alloy LD10.
The two propellant tanks and the inter-tank section form a thrust and weight bearing load structure, with the second stage engines directly mounted on the rear of the fuel tank. The instrument compartment located at the front end of the stage houses the flight control system, navigation platform, and gas canisters.
The second stage engine consists of a main motor with fixed nozzle and a swivelling vernier motor consisting of four small chamber motors.
The second stage is also fitted with four small solid fuel rocket motors, which are fired for only half a second at 3.1 seconds after the stage engine shuts, which reduces the velocity of the rocket stage by 1—1.5 m/s to allow it to separate from the payload.
Secondary Payload Adaptor
A secondary payload can be placed at the front end of the second stage between the instrument compartment and the prime payload fairing. The payload is mounted on a two-piece payload adaptor 2.2 m in diameter, with the upper part of the adaptor providing electric and mechanical interfaces for the prime payload, and the lower part secured to the instrument compartment of the second stage. The two parts are connected by four explosive bolts. After the prime payload is separated from the rocket stage, the booster can adjust its flight trajectory and then detonate the explosive bolts to release the secondary payload into its planned orbit.
On 6 October 1992, a CZ-2C launcher taking off from Jiuquan carried a FSW satellite as its prime payload and the Swedish scientific satellite Freja as secondary payload. The rocket first placed the FSW satellite into a 210 km x 329 km, 63° inclination orbit, before placing Freja into a 600 km x 1,725 km orbit of same inclination.
A spin-stabilised upper stage known as SM has been developed for the CZ-2C to place payload into orbits beyond the reach of the rocket. The stage is powered by a 10.78 kN thrust solid rocket motor burning HTPB/Hydrazine propellant and reaction control system (RCS), enabling the rocket to place 1,250 kg payload into the Geostationary Transfer Orbit (GTO).
A new three-axis stabilised upper stage known as SMA can enable the rocket place 1,900 kg payload into a 600 km Sun Synchronous Orbit (SSO).
The CZ-2C is fitted with a two-piece payload fairing to protect the satellite from aerodynamic forces during flight. The fairing connected together by 12 explosive bolts and then secured to the rocket’s second stage by 8 explosive bolts. The fairing is jettisoned at an altitude of about 120 km.
The CZ-2C has two types of payload fairing available:
Type A fairing is simply a two-piece cone which also serves as payload adaptor. It is installed to the launch vehicle together with the payload on the launch pad.
Type B fairing is a more sophisticated design consisting of nose dome, forward cone section, and honeycombed cylindrical section. Cork panels are installed on the forward cone and cylindrical sections for heat-proof insulation, damp absorption, and noise reduction. The fairing is integrated with the payload adaptor and the satellite inside a ‘clean room’, before they are transported as a whole to the launch pad to be integrated with the launch vehicle.
To accommodate a wide range of loads, the CZ-2C has three adapters that are geared to different loads: Type 937, Type 1194 and Type 1497.
The first stage of the CZ-2C is powered by the YF-21 engine, which consists of a cluster of four parallel YF-20 rocket motors arranged symmetrically at an angle of 2°50’ to the axis of the launch vehicle. Each engine has a swinging nozzle that can be pivoted up to +/-10° at radial direction to provide directional thrust and steering.
The second stage is powered by an YF-22 main motor with fixed nozzle, and a swivelling vernier motor consisting of four YF-23 chambers. The swivelling vernier motor was designed for steering and sustaining propulsion for a further 190 seconds after the shutting of the main motor, thus increasing the payload capacity of the launch vehicle.
Typical Flight Profile
T-3 sec: 1st stage ignition T+0 sec: Liftoff T+8 sec: Pitch-over manoeuvre T+44 sec: LV reaches Mach 1 T+127 sec: 2nd stage ignition T+128.5 sec: 1st stage shutdown T+129 sec: 1st stage jettison T+237.5 sec: 2nd stage main motor shutdown T+250 sec: Payload fairing jettison T+475 sec: 2nd stage vernier motor shutdown T+478 sec: Payload separation
Launch Vehicle Overall Length....31.17 m Body Diameter.....3.35 m Take-off Mass.....192 t Take-off Thrust...2,786 kN T/W-Weight Ratio..1.48 LEO Capacity......2,400 kg First Stage Length............23.72 m Diameter..........3.35 m Gross Mass........151 t Empty Mass........8.60 t Propellant Mass...143 t Engine............YF-21 Propellant........N2O4/UDMH Thrust............2,786 kN (sea-level) Isp...............2,540 N.s/kg (sea-level) Burn Time.........130 s Second Stage Length............8.71 m Diameter..........3.35 m Gross Mass........38.2 t Empty Mass........3.2 t Propellant Mass...35 t Engine (Main).....YF-22 Engine (Vernier)...4 x YF-23 Propellant........N2O4/UDMH Thrust (Main).....720 kN Isp (Main)........2,834 N.s/kg Burn Time (Main)..112 sec Thrust (Vernier)..4 x 46 kN Isp (Vernier).....2,762 N.s/kg Burn Time (Vernier)...287 sec