The liquid propellant rocket engine combustion chamber represents one of the most difficult engineering flow systems in operation. The ignition of the propellants injected into a rocket combustion chamber and the subsequent propagation and anchoring of the flame is an important design consideration for all types of rockets. Reliable ignition has to be guaranteed and the initiated turbulent diffusion flame has to stabilize without over pressure or blow out. The control of ignition in a rocket engine is a critical problem for combustion chamber design. Delayed ignition may lead to high unsteady chamber pressure that can damage the engine (strong ignition) whereas early ignition may not sustain to reach steady state. Predictivity of the models and numerical tools to analyse the ignition transient has still limitations. In H2/O2 rocket combustors, the injected propellants are ignited by a stream of hot gas originating from the igniter device. The hot gas has to mix with the injected propellants to initiate combustion in a situation which is characterized by strong spatial inhomogeneties. This paper presents the numerical study of ignition characteristics of H2 & O2 propellant combination in cryogenic combustor. The main objective is to get an insight into the main processes involved in the ignition of cryogenic engines. The pressure, temperature, velocity profile and propellant mass fraction variation along the combustion chamber is addressed. The characteristics of diffusion flame, shear layer combustion, recirculation zone and flame propagation are also addressed. The reaction mechanism is studied using Eddy Dissipation Model/Finite Rate Chemistry. Numerical simulation results were compared with the experimental data.