Name: Chang Zheng-2 (CZ-2), or Long March 2 Type: Liquid-propellant orbital launch vehicle Contractor: China Academy of Launch Vehicle Technology (CALT, 1st Academy); Shanghai Academy of Spaceflight Technology (SAST, 8th Academy) First Launch: 5 November 1974 Launch site: Jiuquan, Taiyuan, Xichang Status: Operational
The Long March 2 (Chang Zheng 2, or CZ-2) is a liquid-fuelled, two-stage space launch vehicle designed to loft satellites to Low Earth Orbit (LEO). Developed from the Dong Feng 5 (DF-5, CSS-4) intercontinental ballistic missile (ICBM), the CZ-2 also served as the baseline model for the subsequent CZ-2E/F, CZ-3, and CZ-4 launch vehicle families.
In 1965, the Chinese military proposed the DF-5, a two-stage, liquid-propellant ICBM with a range of 10,000 km. The development of the missile was assigned to the 1st Academy of the Ministry of Astronautics, with two U.S.-trained rocket engineers, Tu Shou’e and Ren Xinmin, in charge of the missile’s overall design and rocket engine respectively.
The DF-5 featured a range of new technologies not seen on previous Chinese ballistic missile designs. The first-stage of the missile would be powered by four parallel 75 t-thrust YF-20 engines. The second-stage would employ a 65 t-thrust YF-22 main engine and a vernier engine. The latter was designed for steering and sustaining propulsion after the second-stage main engine cut-off, in order to enable a wide aiming arc for the missile’s warhead in the upper atmosphere. Both stages of the missile would burn a new liquid bipropellant with unsymmetrical dimethylhydrazine (UDMH) as fuel and nitrogen tetroxide (N2H4) as an oxidiser.
In 1967, the 1st Academy was tasked with developing the DF-5 into a space launch vehicle to loft the 1,800 kg FSW (Fanhui Shi Weixing) recoverable satellite into a 180 km Low-Earth Orbit (LEO). So instead of boosting a 3,000 kg mass thermal nuclear warhead into a ballistic suborbital trajectory, the rocket now needed to insert a lighter satellite into Earth orbit, which requires a higher specific impulse (Isp), i.e. lower thrust that sustains for longer periods of time. In response CALT proposed two options: the first option featured a modified second-stage using a 55 t-thrust YF-25 engine, while the second option proposed to increase the rocket’s orbital payload capacity by optimising the flight profile of its second-stage.
As the new YF-25 would not be ready in the near future, the second option was adopted. The 1st Academy initially proposed to delay the ignition of the second-stage after the first-stage jettison, so that the rocket would fly powerlessly for some time before the second-stage ignition. However, this would require a separate jettison device for the first-stage, thus increasing the rocket’s overall structural weight. This proposal could only increase the rocket’s payload capacity by 100 kg. Instead, a second proposal was adopted to keep the vernier thrusters on the second-stage burning for a further 190 seconds after the main engine cut-off. This would increase the rocket’s payload capacity by 500 to 800 kg, at the expense of a longer flight distance to reach the intended orbit.
The DF-5 made its maiden flight on 10 September 1971, but the test was only partially successful. The second test launch on 8 April 1973 ended up with the missile exploding in the mid-air 43 seconds after lift-off. Chinese Premier Zhou Enlai then ordered the DF-5 development to be suspended, and the remaining four test missiles in the same batch to be converted into launch vehicles under the designation Chang Zheng 2 (Long March 2) for the FSW launch missions.
On 5 November 1974, a CZ-2 carrying the first FSW (or Jian Bing 1 in its military designation) recoverable satellite lifted off from the Jiuquan Satellite Launch Centre (JSLC). However, the rocket exploded in the mid-air only 20 seconds into the flight. Subsequent investigations revealed that the failure was caused by a disconnected cable for the pitch rate gyro signal in the rocket’s guidance system. On 26 November 1975, a second CZ-2 rocket was launched from Jiuquan, successfully lofting an FSW satellite into the intended 185 km Earth orbit. This was followed by further two successful launches in December 1976 and January 1978.
In 1979, the PLA ordered six more CZ-2 launchers for the follow-up FSW missions. As the original four CZ-2 vehicles modified from DF-5 test missiles were all used up, the new launchers would be converted from the DF-5’s Batch-02 productions. CALT also used this opportunity to introduce some modifications to the launch vehicle to improve its performance and reliability. To distinguish from the original CZ-2 vehicles, the improved variant was designated CZ-2C (the A and B designations had already been allocated to the three-stage launch vehicle under development, which eventually became the CZ-3 and CZ-4).
In February 1982, the FSW satellite and CZ-2C launch vehicle were approved by the military and Ministry of Astronautics for batch production. 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 mission in August 1987 also carried two French microgravity experiment packages on its piggyback – the first ever Chinese orbital launch for a foreign customer.
In the mid-1980s, to support the launch of the heavier FSW-1 (Jian Bing 1A) satellite, CALT made further improvements to the CZ-2C design, including the addition of a telemetry system. 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.
On 6 October 1992, a CZ-2C 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 x 329 km, 63° inclination orbit, before placing Freja into a 600 x 1,725 km orbit of same inclination.
In 1993, the Chinese space industry entered a contract with the U.S. telecommunication company Motorola to use the CZ-2C to loft 12 satellites for the Iridium global wireless communications satellite network. To support the launch missions, CALT developed a three-axis stabilised, solid rocket upper stage known as “Smart Dispenser” (SD). The upper stage consisted of a 742 kN thrust main motor and four 74.1 kN venire motors, and had its own guidance system to insert two satellites into orbit in a single launch. The vehicle’s second-stage featured an improved engine with higher expansion ratio nozzles, and was stretched to accommodate additional propellants, increasing the payload capacity to LEO from 2,400 kg to 3,000 kg. A new 3.35 m diameter payload fairing was introduced to accommodate dual satellites. Between 1997 and 1999, the CZ-2C/SD made 7 flights from the Taiyuan Satellite Launch Centre (TSLC), orbiting two mock satellites and 12 operational Iridium satellites.
In the early 2000s, CALT made further improvements to the CZ-2C 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 carry additional propellants. The rocket made three successful flights between 2004 and 2006, orbiting two FSW-4 satellites and a Shijian 8 microgravity experiment satellite.
To support the launch of the “Double Star” scientific research satellites, the CZ-2C was added with a spin-stabilised solid-motor upper stage designated SM. The CZ-2C/SM made its first flight from XSLC on 30 December 2003, followed by the second flight from TSLC on 25 July 2004.
The CZ-2C/SMA featuring a new 3-axis stabilised solid-motor upper stage was launched on 6 September 2008, placing two environment monitor satellites, Huanjing 1A and 1B, into their intended orbits.
In the late 1980s, the military proposed the more advanced FSW-2 (Jian Bing 1B) with an increased mass of 2,600 kg. This would exceed the maximum payload capacity of the CZ-2C, and a heavier launch vehicle was required as a result. CALT proposed a redesigned CZ-2C with increased payload capacity, but was rejected by the military due to the development cost. The Ministry of Astronautics decided in 1990 to adopt the CZ-2D proposed by Shanghai Academy of Spaceflight Technology (SAST) for the launch of the FSW-2.
Despite inheriting the CZ-2 designation, the CZ-2D does not have a direct technological lineage to the other variants of the CZ-2 family. The vehicle was developed by the Shanghai-based 8th Academy (Shanghai Academy of Spaceflight Technology, or SAST) based on its Feng Bao-1 (FB-1) and Chang Zheng-4 (CZ-4), which were also derived from the DF-5 design. The CZ-2D is essentially a two-stage version of the CZ-4, with a payload capacity of 3,100 kg to LEO.
The CZ-2D development began in February 1990 and took two years to complete. The first flight of the launch vehicle took place on 9 August 1992, successfully lofting FSW-2-01 into orbit.
The core vehicle consists of two stages connected by an inter-stage structure, all 3.35 m in diameter. Each stage has two propellant tanks: an oxidiser tank at the front and a fuel tank at the rear, connected by an inter-tank ring section. Oxidiser is pumped to the engines via a pipe penetrating through the centre of the rear fuel tank. The two propellant tanks and the inter-tank ring section form part of the vehicle’s thrust and weight bearing load structure and are constructed from high-strength aluminium-alloy LD10. The four strap-on boosters used by the E and F variants are 2.25 m in diameter and have a similar structure to the first-stage. A two-piece payload fairing protects the satellite from aerodynamic forces during flight. The fairing is connected together by 12 explosive bolts and secured to the second-stage by 8 explosive bolts, and is jettisoned at an altitude of about 120 km.
The first-stage comprises (from front to rear): oxidiser tank, inter-tank ring section, fuel tank, engine frame, and tail section. The front end of the forward oxidiser tank is protected by a fibreglass heat insulation layer to prevent damage from the high pressure and hot stream of engine exhaust from the second-stage engine during stage separation. The four main engine motors 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 stage is powered by the YF-21 engine, which consists of a cluster of four parallel YF-20 single-chamber motors arranged symmetrically at an angle of 2°50’ to the axis of the core vehicle. Each motor has a swinging nozzle that can be pivoted up to +/-10° at radial direction to provide directional thrust and steering.
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 is also incorporated with 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. The erected vehicle is supported by four weight bearing points on the tail section, secured to the launch pad’s base unit with explosive bolts, which are detonated less than a second before the vehicle lifts off.
The inter-stage structure comprises an inter-stage shroud and a grid structure. The inter-stage shroud, 3.35 m diameter and 3.2 m in length, houses second-stage engines to protect them from aerodynamic forces, and is connected to the rear of the second-stage’s 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, thus eliminating the need for jettisoning devices to provide the separation impulse and avoiding 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 ring section, and fuel tank. The engines are directly mounted on the rear of the fuel tank. The instrument compartment at the front houses the flight control system, navigation platform, and gas canisters. The stage is powered by an YF-22 main engine with fixed nozzle, and a swivelling vernier engine consisting of four YF-23 chamber motors. The swivelling vernier engine is designed for steering and sustaining propulsion for a further 190 seconds after the main engine cut-off. The stage also has four small solid fuel rocket motors, which are fired for only half a second after the second-stage engine cut-off. This reduces the velocity of the rocket stage to allow the payload to be separated from the stage.
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.
Stages:...................2 Strap-on boosters:........0 Overall length (m):.......31.17 (CZ-2C) 40 (CZ-2C/SD) Take-off mass (t):........192 (CZ-2C), 213 (CZ-2C/SD) Take-off thrust (kN):.....2,786 Thrust-weight ratio:......1.48 LEO capacity (kg):........2,400 (CZ-2C); 3,000 (CZ-2C/SD) Stages 1st-stage 2nd-stage ---------------------------------------------------- Length (m):...............23.72 8.71 Diameter (m):.............3.35 3.35 Gross mass (t):...........151 38.2 Empty mass (t):...........8.60 3.2 Propellant mass (t):......143 35 (CZ-2C) 55 (CZ-2C/SD) Engine:...................YF-21 YF-22 (Main) 4x YF-23 (Vernier) Propellant:...............N2O4/UDMH N2O4/UDMH Thrust (kN)*..............2,786 720 (Main) 4x 46.0 (Vernier) Isp (N.s/kg )*............2,540 2,834 (Main) 2,762 (Vernier) Burn time (sec):..........130 112 (Main) 287 (Vernier) Payload fairing Type A Type B ----------------------------------------------------- Length (m):...............3.144 7.125 Diameter (m):.............2.2 3.35
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