A large complex manifold of an injector allows and

high flow velocities; poor distribution of flow over the chamber cross-section

low flow velocities; poor distribution of flow over the chamber cross-section

high flow velocities; good distribution of flow over the chamber cross-section

low flow velocities; good distribution of flow over the chamber cross-section

Correct Answer : low flow velocities; good distribution of flow over the chamber cross-section

Explanation : For larger manifolds, the injector flow will be lesser and the distribution of flow over the chamber cross-section will be uniform. This can be seen by considering two passages of varying cross-sections and constant mass flow rate of propellants through it.

Recently Updated in Rocket Propulsion Questions

Determine the nozzle throat area for a rocket engine producing a total thrust of 10 kN, chamber pressure of 5.2 MPa and thrust coefficient of 1.9

10.12 cm²

11.20 cm²

12.10 cm²

21.10 cm²

Correct Answer : Option (A) : 10.12 cm²

Explaination : Nozzle throat area A_t = F/P₁C_F, where P₁ is the chamber pressure, C_F is the thrust coefficient and F is the total thrust provided by the rocket engine.
Then A_t = 10000/(5.2 x 10⁶ x 1.9)
= 10.12 cm².

The pressure drop across injector is usually set between _________ of the chamber pressure.

50 and 70%

45 and 65%

30 and 45%

15 and 25%

Correct Answer : Option (D) : 15 and 25%

Explaination : The pressure drop across the injector is set between 15 and 25% of the chamber pressure. This is done partly for obtaining high velocities of propellant injection. High velocities will help in easy atomization and droplet breakup which in turn helps in complete combustion.

Calculate the flow rate of LOX to the combustion chamber in a LOX-H2 rocket engine if the total propellant flow rate is 17 kg/s and mixture ratio is 3.2.

12.95 kg/s

13.67 kg/s

15.92 kg/s

17.81 kg/s

Correct Answer : Option (A) : 12.95 kg/s

Explaination : m?_o = m?_r/(r+1)
So m?_o = 17*3.2/4.2 = 12.95 kg/s.