Launch mass is the satellite fully fuelled at lift-off; dry mass is the satellite without its propellant — often less than half the launch mass.
When ISRO says a satellite weighs 1,425 kg, that is usually its lift-off mass: everything, including the propellant in its tanks. The dry mass is what is left once that propellant is taken away — the structure, electronics, panels and payload.
The difference is the satellite's fuel for its whole life.
Worked example: Kalpana-1 carried 1,060 − 498 = 562 kg of propellant, about 53 percent of its launch mass. The IRNSS satellites carry 1,425 − 600 = 825 kg, about 58 percent. More than half of each satellite, at launch, was propellant.
Why so much? Rockets such as PSLV and GSLV often leave satellites in a stretched transfer orbit; NVS-02, for example, was placed in one by GSLV-F15. The satellite then burns its own propellant to climb to its final orbit, and keeps some to stay there for years. Kalpana-1's launch, PSLV-C4 on 12 September 2002, was PSLV's seventh flight and the first time it put a satellite into this kind of transfer orbit.
A rocket's capacity is quoted for the launch mass. So every kilogram of propellant a satellite needs is a kilogram that cannot be payload. That is the main reason ISRO is moving towards electric propulsion: it says the mass saved could allow more transponders on communication satellites. In March 2025 it finished a 1,000-hour test of a 300 mN xenon plasma thruster, which it plans to fly on its Technology Demonstration Satellite, TDS-01.
GSAT-19, at 3,136 kg, was the heaviest satellite launched from Indian soil when it flew in June 2017.
Check the altitude of any NavIC or GSAT satellite: most of the propellant counted in its launch mass was spent climbing from a low perigee to that height. Open the ISRO satellite tracker
The link between the two masses is the rocket equation:
Δv = v_e × ln(m_launch ÷ m_dry)
It tells you how much total velocity change the satellite can make on its own propellant.
Kalpana-1 (ISRO masses), assuming a bipropellant exhaust speed of v_e ≈ 2,940 m/s (Isp ≈ 300 s, within NASA's 150–310 s band for chemical mono- and bipropellants):
m_launch ÷ m_dry = 1,060 ÷ 498 = 2.13, so ln 2.13 = 0.755
Δv ≈ 2,940 × 0.755 ≈ 2.2 km/s
IRNSS:
1,425 ÷ 600 = 2.375, ln 2.375 = 0.865
Δv ≈ 2,940 × 0.865 ≈ 2.5 km/s
Propellant fraction:
This is the budget for climbing out of the transfer orbit, correcting inclination and holding station for the design life. Because Δv depends on the logarithm of the mass ratio, each extra km/s costs more propellant than the last: to double Δv you must square the mass ratio.
With electric propulsion at v_e ≈ 17,600 m/s (Isp ≈ 1,800 s), the same 2.5 km/s needs a mass ratio of only e^(2,500 ÷ 17,600) = e^0.142 ≈ 1.15 — about 13 percent propellant instead of 58.
1,060 − 498 = 562 kg, more than half its launch mass.
Rockets usually leave them in a transfer orbit, so they must burn their own propellant to reach geostationary orbit and then to keep station for years.
The mass of a spacecraft without propellant — its structure, systems and payload.