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LS3: Combustion & Propellants

contactOriginal Author: Matt Morningstar:  Matt Morningstar '21, matt_m@mit.edu

Lecture Zoom Recording

Specific Impulse and the Rocket Equation, Revisited

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As we discussed in the first section, a high combustion temperature is most efficient. So you might expect that rocket engines use a stoichiometric mixture ratio. Unfortunately, the combustion temperatures achieved at a stoichiometric mixture ratio are so hot that they exceed the limits of our material/cooling capabilities. As a result, nearly all rocket engines use a fuel-rich mixture ratio (ox-rich combustion presents a number of issues, which is why an ox-rich mixture ratio is not commonly used in the main combustion chamber). It is also true that a stoichiometric mixture ratio (even if we could withstand it's temperature) does not always yield maximum Isp. In the case of hydrogen/oxygen, it is more optimal to operate fuel-rich, as it leaves more light H2 molecules in the exhaust.

Questions:

Starting Parameters

The following parameters are given. Obviously, in real engine design, these (very crucial) parameters won’t be handed to you, and the determination of their optimal values for the requirements of the engine will be a very involved process.

  • Propellants

    1. Fuel: Methane (CH4)

    2. Oxidizer: LOX (O2)

    3. Mixture ratio = O/F = 3

  • Chamber pressure (4 MPa)

  • Desired thrust = 3500 N

  • Operating at sea level

 

    • Is the given mixture ratio fuel-rich or ox-rich? How do you know? What does this mean?

    • What is the Tc for this flow (combustion temperature, aka stagnation temperature, aka adiabatic flame temperature)?  Would increasing the mixture ratio lead to a higher or lower combustion temperature?                             

    • Why might we not want to use the stoichiometric mixture ratio?

    • For this combination of propellants, what is the value of:

      •  ratio of specific heats (gamma)?

      •  the specific gas constant (R)? 

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2. Exhaust Velocity

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the

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exhaust

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What is the required mass flow rate to achieve the thrust target?

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What is the required throat area?

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What is the Isp of this engine?

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