Ignite the Future Workshop

thrust, propulsion, propellant combustion - burning oxidizer , fuel

Things to learn

• back pressure
• understanding different pressure metrics (bar, atm, pascals) and tank pressures, chamber pressure
• Intercollegiate Rocket Engineering Competition - Space port america cup
•  Experimental Sounding Rocket Association (ESRA)
• how does a hand pump work
• what are turbines and how do they work , same concept of dynamo
• autogenous pressurisation
• how does throttling work in a liquid rocket engine (raptor and merlin), throttlability
• specific heat capacity in thermodynamics (cooling and heating) $T_0$ (current temperature) to $\Delta T$ (the required change in temperature)
• flame quenching ,
• stability margin of the rocket
• injectors, pintle injector
• mixing of fuel and oxidizer properly for the combustion and thrust

wifi credentials: shaastra002 , Chennai@36*

potassium nitrate - $KN{O}_3$ - oxidizer for the solid propellants

bipropellant - fuel and oxidizer monopropellant - decombustion, a single chemical substance that acts both as a fuel and oxidizer

• $N_{2}O$ - monopropellant liquid
• $AP$ - solid mono-propellant

tccc - Thrust chamber cooling channel (around the engine bell)

why do gases and liquids flow - difference in potential energy for liquids and different in pressure for gases

$\rho gh$ - equation for flow

to stop the flow, you have a mechanical obstruction that can hold the pressure, or you equalize the pressure

equillibrium is a macroscopic phenomenon, dynamic equllibrium is a microscopic phenomenon

to avoid back flow of propellant we make use of check valve

back flow leads to RUD

liquid prop equation (?? unsure - re verify) $A{C}_P=\sqrt{DP\rho }$

700kN - commercial rocket engines thrust 5 bars to 250 bars 15 Kg per second

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Solid rocket motors

• grain design - star (higher suface area), changes to a circle later on, which reduces the overall thrust sent
• disadvantages - solid motors have a problem of blow holes (the propellant is not smooth when casting - air traps)
• solid rocket propellant is completely dependent on the chamber pressure

aluminium 6061 - outer casing for the solid rocket motor of abhyuday

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Liquid rockets

• for thrust vectoring, due to this the chambers of liquid rocket engines should be least heavy

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$d^{2}law$ - for hybrid rockets combustion efficiency - the entire fuel that’s injecting into the chamber is not combusting

why do we need a pre combustion chamber in hybrid rocket motors

parabin fuel - intermediate droplet when the wax melts, entrainment droplets. residence time- the average time in which the combustion is present in the chamber in liquids - the minimum nozzle size comes from the residense time, and a few other parameters a port for pressure transducer isentropic style of combustion - theoritical and not in experimental - happens in the nozzle pressure loss decreases the efficiency

F = mass flow rate * velocity , $\dot{m}\ast Isp$

non-dimensional numbers, the ratios between the l / d - 1.53 ratio - 90 mm / 40 mm length / diameter of the ratio of the post combustion chamber

insulation material between the outer casing and the combustion chamber - HYLAM the black carbon is formed by the HYLAM material

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calculation

NASA-CEA analysis → gives us the Isp and C* NASA-SP125

500 N of thrust is required

$F=\dot{m}\ast Isp$

$F=\dot{m}\ast Isp\ast {\eta }_{Isp}$ here ${\eta }_{Isp}$ is the efficiency term

$\dot{m}={\dot{m}}_{Fuel}+{\dot{m}}_{Oxidizer}$

500 = 0.2717391304

stoichometry ratios O/F ratio - oxidizer flow rate / fuel flow rate

2.4 ratio for O/F for maximum thrust for parabin

regression rate - the change in radial length per unit time

$\dot{r}=\frac{r_f-r_i}{t}$

$\dot{r}=a\ast G_{ox}^n$

Gox range = > mass flow rate of oxygen / Ap (fuel port area ) $G_{ox}range=\frac{\dot{m}}{A_p}$

the combustion was chuffing - combustion instability problem

45 Kg / m2 ⇐ G0x ⇐ 800 kg / m2 ⇒ parafinic for Gox

density = 900 kg / m2

diameter = 17.4mm

2.7 seconds

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choke flow equation m dot = Pc * A p / C*

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