S. Ramakrishnan, R. N. Tyagi, S. S.
Balakrishnan* and Sojan Thomas
The fifth Indian Polar Satellite Launch
Vehicle mission, PSLV-C2, which took off on 26 May 1999, at 1152 h IST, marks an
important milestone in the Indian Space Programme. This is the first time that an Indian
launch vehicle carried multiple satellites, one primary satellite IRS-P4, a 1036 kg
Indian remote sensing satellite (IRS-P4) or OCEANSAT-1; and two auxiliary payloads into
720 km polar Sun-Synchronous Orbit (SSO).
The two auxiliary payloads, or
microsatellites were the KITSAT-3 weighing about 107 kg developed by the Satellite
Technology Research Centre, Korea Advanced Institute of Science & Technology, Republic
of Korea; and the DLR-TUBSAT weighing 45 kg developed by Technical University of
Berlin (TUB) and German Space Centre (DLR). With the successful launch of these
microsatellites by ISRO, its entry into the multi-million commercial launch services
market has been signalled.
This article gives the details on the
PSLV-C2 mission; highlighting the studies, changes, and preparation made with reference to
launch of auxiliary satellites along with the main satellite.
PSLV-C2 mission: Flight no. 5
PSLV is a four-stage vehicle, that employs solid and liquid stages
alternately (Figure 1). After three developmental missions, the fourth mission of PSLV
(PSLV-C1) placed the operational IRS-1D satellite into polar orbit during September 1997.
Incidentally, PSLV-C1 was the first launch in PSLV Continuation Programme, under which the
Government had approved the production of 9 vehicles.
Changes from C1 to C2 mission
This section provides the modifications carried out on the PSLV equipment bay
to accommodate the auxiliary satellites, development of the satellite separation system,
structural tests to analyse the adequacy of the satellite structure, and on the mission
analyses performed to ensure collision-free separation even under worst conditions.
Accommodation of auxiliary satellites
PSLV configuration used for the C2 mission was the same as that was flown in
the previous mission. But to accommodate the auxiliary payloads, provision was made on the
vehicle equipment bay (VEB), on diametrically opposite sides to carry up to two
microsatellites (< 100 kg, 600 ´ 600 mm
size, and 800 mm ht.) in a piggy-back mode. This will be used as a standard
arrangement for all the future flights, in order to enable piggy-back ride for this class
of satellites, for scientific and technology-providing missions (Figure 2).
Separation system for auxiliary
To mount the auxiliary satellites on the VEB and to separate them, ball-lock
separation systems, 358 mm dia for KITSAT-3 and 230 mm dia for DLR-TUBSAT, were
developed and tested. The ball-lock mechanism consisted of two rings, the inner ring
attached to the bottom surface of the satellite and the outer ring attached to the VEB.
These two rings were held together by a number of hardened steel balls which locked the
inner race to the outer ring. Redundant pyrothrusters rotate the outer ring by about 4° to cause the
radial escape of the balls through aligned holes, thereby releasing the inner ring
attached to the satellite. The separation velocity to the auxiliary payload was provided
by the energy stored in the compressed springs, which were located symmetrically at the
interface. A nominal velocity of 1.0 m/s was imparted to the microsatellites.
Structural model test
The above configuration changes, incorporated for the microsatellites,
required validation of these changes. For this, the fourth stage along with mock-up
equipment bay and the structural models of KITSAT-3 and DLR-TUBSAT were stacked, in order
to check for their mechanical and electrical interfaces compatibility, and were vibrated
(both sine and random vibrations) in order to determine the characteristics along
longitudinal and lateral axes of the spacecraft. This test also provided input for coupled
load analysis and responses at critical locations on the spacecraft.
The vibrated stack was tested for separation of satellites, and also for the
measurement of spacecraft separation velocities and shock levels at the spacecraft
interface and on the vehicle equipment bay.
The mission analysis performed the following:
load analysis for loads and frequency adequacy of the satellite structures.
dynamic separation analysis for worst case conditions without collision with the vehicle
rotational velocities imparted to the separating satellite(s) under worst case conditions.
propagation of relative motion among the various satellites: OCEANSAT; KITSAT-3; and
A detailed Monte-Carlo analysis was conducted considering dispersion values
for various parameters like CG offsets, MI, mass, clearance between satellite face and the
4th-stage tankage, differential spring velocities, etc. and a positive design margin was
A brief summary of the characteristics of PSLV stages is given in Table 1.
PSLV-C2: Flight 05 trajectory
The trajectory design features of PSLV-C2 are such that they allow the
launcher to ascend vertically up to T+5 s. After this, the vehicle is supposed to
roll by 5° to orient
itself towards 140° Azimuth with
reference to true north, and is supposed to pitch down as per the pre-determined pitch
program stored on-board.
At the end of I-stage burn and ignition of the 2nd stage, the vehicle would
be maneuvered in yaw plane also in order to orient the vehicle in such a way that it would
reach the mission inclination upon injection. During the II-stage burn, at
T + 162 s, the heat shield would be jettisoned at an altitude
> 115 km. After this, the vehicle closed loop guidance (CLG) system would
steer the vehicle till the injection of the satellite(s). To cater to performance
deviations of the lower three stages, sufficient guidance margin, by way of additional
fuel, has been provided in the IV-stage and, thus, CLG would ensure the right orbit as
long as the vehicle dispersions are within 3s
Soon after the guidance cut-off of the 4th stage, with the achievement of the
mission target parameters, the IRS-P4 would get separated followed by the separation of
auxiliary satellites KITSAT-3 and DLR-TUBSAT following pre-fixed yaw maneuvers to avoid
collision. The sequence of separation is depicted in Figure 3.
Based on the injection state available from the on-board inertial navigation
system, the down-range radar and the s-band telemetry tracking, the preliminary orbits
have been determined for further acquisition of satellites by the respective satellite
agencies. The total mission from take-off would be about 1150 s.
The values of predicted and actual orbit obtained have been listed in Table
IRS-P4 (OCEANSAT) satellite
Indian remote sensing satellite, IRS-P4 is the eighth in the series and the
first one dedicated for ocean studies. It carried an ocean colour monitor (OCM) and multi-
frequency scanning microwave radiometer (MSMR) for ocean-related applications, and hence
was named OCEANSAT-1. The stowed view of IRS-P4 is given in Figure 4.
The mission objectives
· To gather
data for oceanographic, land (vegetation dynamics) and atmospheric applications.
· To provide
new application areas using data as complementary/supplementary to the already operating
remote sensing satellites.
· To provide
opportunity for conducting technological/scientific experiments that are of relevance to
Mass at lift-off (kg)
of 2.8 ´ 2.0 ´ 2.6
Length when fully
Attitude and orbit
After launch, the satellite has been under the control of ISRO telemetry,
tracking and command network (ISTRAC) at Bangalore and the payload data is being received
at the National Remote Sensing Agency (NRSA), Hyderabad.
KITSAT-3 is an engineering test satellite whose primary mission objective has
been to develop various fundamental technologies for high performance microsatellites to
qualify them in the low-earth-orbit space environment
It was developed by Satellite Technology Research Center (SaTReC), Korea and
the Advanced Institute of Science and Technology (KAIST), Republic of Korea. The view of
the fully assembled KITSAT-3 is shown in Figure 5 and the exploded view in Figure 6.
Summary of major features
The major features of KITSAT-3 are summarized as follows:
: < 110 kg
495 ´ 630 ´ 854(H) mm
Power : 150 W max.
fixed and 2 deployable solar panels)
stabilized attitude control
accuracy < 0.5°)
sensors, sun sensor, IR earth horizon sensor,
fiber optic gyros
kbps in S-band
Mbps in X-band
Gbits flash memory
earth imaging system (MEIS)
environment scientific experiment (SENSE)
Multi-spectral earth imaging system
The MEIS consists of a catadioptic telescope, prism blocks, focal plane
assembly, video signal processing electronics, camera flight processors (CFP), and
solid-state mass memory units. The MEIS is of push-broom type and produces images of the
earths surface in three spectral bands. Its main characteristics are:
6.5 kg (3 kg for telescope)
· Power: 15
Three linear CCDs with 3456 pixels
19.1 µrad (GSD ~ 15 m)
· FOV: 3.8° (swath width
~ 50 km)
bands: 520620, 620690, 730900 nm
focal length: 570 mm
· F number:
· MTF: 20%
at nyquist frequency
Space environment scientific
The SENSE consists of the following four sub-systems.
particle telescope (HEPT)
effects on micro-electronics (REME)
temperature probe (ETP)
HEPT is not only being used to measure the particle energy entering the
telescope but also has the capability to identify the particle species. Four silicon
detectors have been used along with blocking materials of aluminium and copper, which are
being used to control the particle energy reaching each detector. Moreover, it can also
measure the pitch angle distribution of particle energies.
REME consists of TDE monitor and SEU monitor units. While the TDE monitor is
being used to measure the long-term accumulated ionizing radiation dose in SiO2
for three locations in the KITSAT-3 bus, the SEU monitor is being used to perform, a
series of testing to measure the SEU characteristics of memory devices.
SMAG is being used to measure the magnetic field using a flux-gate type
sensor and is also being used as a diagnostic tool for global and local geomagnetic
disturbance and current systems, low frequency waves in the ambient environment, and of
the waveparticle interaction. Thus it will provide the information on the magnetic
field direction for HEPT and plasma gyro-frequencies.
ETP is being used to measure the electron temperature in high latitude
regions. It was specifically developed in order to study the occurrence of anomalous
heating phenomenon found in the South Atlantic Anomaly, as well as to find any
relationship it has with other plasma parameters such as energetic particles and plasma
KITSAT-3 was separated from the VEB deck after the separation of the IRS-P4.
After 3.5 h, the telemetry was switched on and the first contact with the command
ground station in Korea was at 7.5 h after the separation. After the spacecraft
health checks were completed, the solar panel was deployed producing full power of
180 W required for operations. The first set of pictures that have been received
reveal the quality of the pictures. These pictures can be viewed at the web site
The microsatellite DLR-TUBSAT was jointly developed by the Technical
University of Berlin (TUB) that was responsible for the satellite and the German Aerospace
Centre (DLR) that was responsible for the payload.
The cube-shaped satellite measures 32 ´ 32 ´ 32 cm
and weighs 45 kg. The task of the satellite is earth observation with high resolution
(better than 10 m ground pixel resolution). The satellite is being operated from the
TUB Satellite Control Centre.
DLR-TUBSAT has a payload module, a power module and an attitude control
module. The exploded view is shown in Figure 7.
The payload module contains two fore field sensors with low- and
medium-resolution telescope and a high-resolution telescope with a focal length of
1 m, a pixel size of 8.3 µm and a ground pixel resolution of 7 m. Each
CCD-chip contains 752 ´ 582
pixel, and each camera can transmit both video images in CCIR-standard and signal digital
pictures. The focal length of the fore field cameras is 16 mm and 50 mm
respectively. The S-band antenna is physically located on the attitude control module in
order not to obstruct the field of view of the payload sensors, and the S-band transmitter
is located close to the antenna. For the transmission of pictures using analog video, a
bandwidth of 8 MHz has been selected and the transmission of single pictures occurs
at 128 kbaud. The beamwidth of the antenna is 70°.
The power module contains the batteries, the power control unit, and the two
UHF transceivers as well as the two UHF antennae. Four duplex NiH2 battery
cells of 12 Ah capacity have been used to support an unregulated 10 V bus that
is charged by four identical solar panels, each containing a single string of 34 silizium
solar cells. The supply of short circuit current to each of the panels is 960 mA. The
UHF transceiver receives and transmits data via FFSK modulation at a rate of 1200 baud.
Each of the transceivers has been connected to one of the UHF antennae. Both the
transceivers are nominally operating in parallel, in a listening mode. As long as no order
is received from the ground, the satellite remains silent.
The attitude control module includes three reaction wheel-gyro pairs as well
as the on board data handling system (OBDH). The OBDH system is being used to transmit
targets (angles or angular rates) to the control loops within the reaction wheels. The
targets can also be set by ground command.
In its standby (hibernation) mode, the satellite tumbles at a natural rate of
typically 0.1 rpm. The acquisition sequence will start with a rate reduction command,
followed by a coarse sun acquisition maneuver, using the information of the solar arrays
and a coarse earth acquisition maneuver using the same (albedo) information. This will be
followed by the TV transmission and the user on the ground can interactively steer via the
wheel-gyro pairs, in a rate-control mode, pointing towards any interesting scene on the
ground. Once satisfied with the view, a photo can be commanded with high, medium (color)
or low resolution. The fully assembled view of the DLR-TUBSAT is shown in Figure 8.
A ball-lock type separation system has been attached to the bottom plate of
the satellite in order to eject the satellite with a separation velocity of approximately
1 m/s. The unit with a pitch circle diameter of 23 cm is supplied by
PSLV injected DLR-TUBSAT, after the separation of both IRS-P4 and KITSAT-3,
about 127 s after cut-off the 4th stage.
During the launch phase however, the NiH2 batteries were
discharged for safety reasons. After separation from the launcher, as per the plan, the
solar panels used to re-charge the batteries within the following 7 h. However,
contrary to the planned schedule, the DLR-TUBSAT was tracked on the 2nd orbit itself and
the telemetry was received over the university ground station in Berlin. Since then, the
satellite has been functioning very well and good pictures are being received from the
With the demonstration of the fourth consecutive successful launch, PSLV has
established its versatility. It can provide multiple launches to Low Earth Orbit (LEO),
polar sun-synchronous missions and also for Geo-synchronous Transfer orbit (GTO) missions.
The payload capability is more than 3000 kg in 400 km LEO, up to 1200 kg polar
at 800 km altitude and 850 kg in GTO. Further planned improvements of PSLV are:
· A high
performance upper stage by this year end which would increase the payload capability
further in all the above orbits.
· A single
engine version of the 4th liquid stage to increase the payload volume inside the heat
shield by about 25%.
· a dual
launch adapter to utilize the above-increased payload volume to enable the launch of a mix
With the above and the PSLV-C2 mission success, PSLV has opened the way for
dedicated commercial launch services for bigger and heavier satellites.