Index

CHAPTER 2

GENERAL DESCRIPTION TO LM-2C

2.1 Summary

The Long March 2C (LM-2C) is a liquid launch vehicle mainly used for Low Earth Orbit (LEO) missions. The LM-2C is most frequently used version of Long March Launch Vehicles which had 14 consecutive successful flights till October of 1994. In order to meet the userís need, China Academy of Launch Vehicle (CALT) developed a new smart dispenser upper stage, the LM-2C/SD has been used commercially in the late 1990s and conducted 7 consecutive successful launches for Iridium program.

The LM-2C launch vehicle now provides two versions:

LM-2C provides flexible mechanical and electrical interfaces and length-adjustable fairing for various SCs. The launch environment impinging on SC, such as vibration, shock, pressure, acoustics, acceleration and thermal environment, meets the common requirements in the commercial launch services market.

LM-2C uses JSLC as its main launch site, it also can be launched from XSLC and TSLC.

2.2 Technical Description

The two configurations of LM-2C share common Stage-1, Stage-2 and fairing. The total length of LM-2C is 42 meters. The diameter of the Stage-1, Stage-2 and fairing is 3.35 meters. The storable propellants of N2O4/UDMH are fueled. The lift-off mass is 233 tons, and lift-off thrust is 2962 kN. Table 2-1 shows the major characteristics of LM-2C.

 

Table 2-1 Technical Parameters of LM-2C

Stage

First Stage

Second Stage

CTS*

Propellant

UDMH/

N2O4

UDMH/

N2O4

HTPB/Hydrazine

Mass of Propellant (kg)

162706

54667

125/50

Engine

DaFY6-2

DaFY20-1(main)

DaFY21-1(verniers)

Solid Motor/

RCS(Reaction Control System)

Thrust (kN)

2961.6

741.4 (main)

11.8´ 4(verniers)

10.78 (solid motor)

Specific Impulse (N· s/kg)

2556.5

(On ground)

2922.37(main)

2834.11(verniers)

(In vacuum)

2804 (solid motor)

Stage Diameters (m)

3.35

3.35

2.7

Stage Length (m)

25.720

7.757

1.5

Note: * CTS is detailed in Paragraph 2.4 of this Chapter.

2.3 LM-2C System Composition

LM-2C consists of rocket structure, propulsion system, control system, telemetry system, tracking and safety system, separation system, etc.

2.3.1 Rocket Structure

The rocket structure functions to withstand the various internal and external loads on the launch vehicle during transportation, hoisting and flight. The rocket structure also combines all sub-systems together. The rocket structure is composed of first stage, second stage and fairing.

The first stage includes inter-stage section, oxidizer tank, inter-tank section, fuel tank, rear transit section, tail section, propellant feeding system, etc. The second stage includes launch vehicle adapter, vehicle equipment bay (VEB), oxidizer tank, inter-tank section, fuel tank, propellant feeding system, and launch vehicle adapter etc. The launch vehicle adapter connects the SC with LM-2C and conveys the loads between them. The international wide-used 937B,1194A adapters are provided. The fairing, with two halves, is composed of dome, forward cone section and cylindrical section. See Figure 2-1 for LM-2C/CTS configuration.

2.3.2 Propulsion System

The propulsion system, including engines and pressurization/feeding system, generates the thrust and control moments for flight. Refer to Figure 2-2a&b.

The first stage and second stage employ storable propellants, i.e. nitrogen tetroxide (N2O4) and unsymmetrical dimethyl hydrazine (UDMH). The propellant tanks are pressurized by the self-generated pressurization systems. There are four engines in parallel attached to the first stage. The four engines can swing in tangential directions. The thrust of each engine is 740.4kN, and the total thrust of first stage engine is 2961.6. There are one main engine and four vernier engines on the second stage. The total thrust is 798.1kN. CTS takes a solid motor as its main engine and Reaction Control System (RCS) for attitude-adjustment. (Detailed in Paragraph 2.4 of this chapter.)

The propulsion system has experienced a lot of flights and its performance is excellent. Figure 2-2a indicates the system schematic diagram of the first stage engines, Figure 2-2b shows the system schematic diagram of the second stage engine.

2.3.3 Control System

The control system is to keep the flight stabilization of launch vehicle and to perform navigation and/or guidance according to the preloaded flight software. The control system consists of guidance unit, attitude control system, sequencer, power distributor, etc. See Figure 2-3a,b&c for the system schematic diagram of the control system. CTS adopts an independent control system. (Detailed in Paragraph 2.4 of this chapter.)

The guidance unit provides movement and attitude data of the LV and controls the flight according to the predetermined trajectory. The attitude control system controls the flight attitude to ensure the flight stabilization and SC injection attitude. For Two-stage LM-2C configuration, the control system re-orient LM-2C following the shut-off of vernier engines on Stage-2. The launch vehicle can spin up the SC according to the requirements from the users. The spinning rate can be up to 10 rpm. The sequencer and power distributor are to supply the electrical energy for control system, to initiate the pyrotechnics and to generate timing signals for some events.

2.3.4 Telemetry System

The telemetry system functions to measure and transmit some parameters of the launch vehicle systems. The telemetry system consists of two segments, on-board system and ground stations. The on-board system includes sensors/converters, intermediate devices, battery, power distributor, transmitter, radio beacon, etc. The ground station is equipped with antenna, modem, recorder and data processor. The telemetry system provides initial injection data and real-time recording to the telemetry data. Totally, about 300 telemetry parameters are available from LM-2C. Refer to Figure 2-4. CTS has its own telemetry system. (Detailed in Paragraph 2.4 of this chapter.)

2.3.5 Tracking and Range Safety System

The tracking and range safety system works along with the ground stations to measure the trajectory dada and final injection parameters. The system also provides range safety assessment. The range safety system works in automatic mode and remote-control mode. The trajectory measurement and range safety control design are integrated together. See Figure 2-5, and refer to Chapter 9.

2.3.6 Separation System

There are three separation events during two-stage LM-2C flight phase, i.e. Stage-1/Stage-2 Separation, Fairing Jettisoning and SC/Stage-2 Separation.

For LM-2C/CTS, there is a SC/CTS separation after SC/CTS stack separates from Stage-2. See Figure 2-6 for LM-2C/CTS separation events.

 

2.4 CTS Introduction

CTS is a three-axis stabilized upper stage compatible with two-stage LM-2C. CTS consists of Spacecraft Adapter and Orbital Maneuver System (OMS). LM-2C/CTS can deliver the spacecrafts into the LEO (h>500 km) or SSO.

LM-2C injects SC/CTS stack into a transfer orbit (Hp=200km, Ha=400~2000km). CTS is ignited at the apogee and enters the target orbit of 400~2000km. CTS re-orients the stack according to the requirements and deploys the spacecrafts. CTS is capable of de-orbiting after spacecraft separation. See Figure 2-7 for typical CTS configuration.

Figure 2-7 Typical CTS Configuration

2.4.1 Spacecraft Adapter

The spacecraft adapter functions to install and deploy the Spacecrafts. LM-2C/CTS provides specific spacecraft adapter according to userís requirements.

2.4.2 Spacecraft Separation System

The separation system can separate the spacecrafts following the insertion to the target orbit. The separation system will be designed to meet the user's requirements on separation velocity, pointing direction and angular rates, etc. The spacecraft are generally bound to the dispenser through low-shock explosive nuts. The separation springs provide the relative velocity. The explosive nuts can be provided by either CALT or SC side.

2.4.3 Orbital Maneuver System

The orbital maneuver system consists of main structure, solid rocket motor (SRM), control system, reaction control system (RCS) and telemetry system.

Diameter

0.54 m

Total Length

<0.9 m

Total Mass

<160 kg

Propellant Mass

121.7 kg

Specific Impulse

2804 m/s

Total Impulse

341.3 kN· s

Burn Time

35 sec.

See Figure 2-3b&c.

2.6 Missions to be Performed by LM-2C

Two-stage LM-2C is a standard LEO launch vehicle with launch capability of 3366 kg (h=200 km, i=63° ). Furnished with suitable upper stages, LM-2C can perform various missions, such as LEO, SSO. Refer to Table 2-3. LM-2C can carry out multiple launches.

Table 2-3 Typical Specification for Various Missions

 

Version

Orbital Requirements

Launch Capacity

Launch Site

LEO

Two-stage LM-2C

Hp=185~400km

Ha=185~2000km

3366 kg (200km/63° )

JSLC

LEO

LM-2C/CTS

Hp=400~2000km

Ha=400~2000km

2800 kg (500km/50° )

JSLC

SSO

LM-2C/CTS

H=400~2000km

1456 kg (900 km)

JSLC

 

 

2.7 Definition of Coordinate Systems and Attitude

The Launch Vehicle (LV) Coordinate System OXYZ origins at the LVís instantaneous mass center, i.e. the integrated mass center of SC/LV combination including adapter, propellants and fairing, etc. if applicable. The OX coincides with the longitudinal axis of the launch vehicle. The OY is perpendicular to axis OX and lies inside the launching plane opposite to the launching azimuth. The OX, OY and OZ form a right-handed orthogonal system.

The flight attitude of the launch vehicle axes is defined in Figure 2-9. Spacecraft manufacturer will define the SC Coordinate System. The relationship or clocking orientation between the LV and SC systems will be determined through the technical coordination for the specific projects.

Figure 2-9 Definition of Coordinate Systems and Flight Attitude

2.8 Spacecraft Launched by LM-2C

Till June 1999, LM-2C has successfully launched 14 recoverable satellites and 12 Iridium satellites listed in Table 2-4.

Table 2-4 Spacecrafts Launched by LM-2C

LV

Date

Payload

SC Manufacturer

Mission

Launch Site

Result

LM-2C F-01

75.11.26

FHW-1

China

LEO

JSLC

Success

LM-2C F-02

76.12.07

FHW-1

China

LEO

JSLC

Success

LM-2C F-03

78.01.26

FHW-1

China

LEO

JSLC

Success

LM-2C F-04

82.09.09

FHW-1

China

LEO

JSLC

Success

LM-2C F-05

83.08.19

FHW-1

China

LEO

JSLC

Success

LM-2C F-06

84.09.12

FHW-1

China

LEO

JSLC

Success

LM-2C F-07

85.10.21

FHW-1

China

LEO

JSLC

Success

LM-2C F-08

86.10.06

FHW-1

China

LEO

JSLC

Success

LM-2C F-09

87.08.05

FHW-1

China

LEO

JSLC

Success

LM-2C F-10

87.09.09

FHE-1A

China

LEO

JSLC

Success

LM-2C F-11

88.08.05

FHW-1A

China

LEO

JSLC

Success

LM-2C F-12

90.10.05

FHW-1A

China

LEO

JSLC

Success

LM-2C F-13

92.10.05

Freja/FHW-1A

Sweden/China

LEO

JSLC

Success

LM-2C F-14

93.10.08

FHW-1A

China

LEO

JSLC

Success

LM-2C F-15

97.09.01

Iridium-DP

Motorola

LEO

TSLC

Success

LM-2C F-16

97.12.08

Iridium-D1

Motorola

LEO

TSLC

Success

LM-2C F-17

98.03.26

Iridium-D2

Motorola

LEO

TSLC

Success

LM-2C F-18

98.05.02

Iridium-D3

Motorola

LEO

TSLC

Success

LM-2C F-19

98.08.20

Iridium-R1

Motorola

LEO

TSLC

Success

LM-2C F-20

98.12.19

Iridium-R2

Motorola

LEO

TSLC

Success

LM-2C F-21

99.06.12

Iridium-R3

Motorola

LEO

TSLC

Success

2.9 Upgrading to LM-2C

Some improvements or upgrading will be taken to make LM-2C launch vehicle more competent or flexible for the market requirements.