Propellant Mass Flow Rate Calculator

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Rocket Propulsion

Components of Gas Turbine

Jet Propulsion

Thermodynamics and Governing Equations

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✖The nozzle throat area refers to the cross-sectional area of the narrowest part of a propulsion nozzle, known as the throat.ⓘ Nozzle Throat Area [A_t]			+10% -10%
✖Inlet Nozzle Pressure represents the pressure of the incoming air or propellant before it enters the combustion chamber or the turbine section.ⓘ Inlet Nozzle Pressure [P₁]			+10% -10%
✖The specific heat ratio describes the ratio of specific heats of a gas at constant pressure to that at constant volume.ⓘ Specific Heat Ratio [γ]			+10% -10%
✖Temperature at Chamber typically refers to the temperature inside a closed chamber or enclosure.ⓘ Temperature at Chamber [T₁]			+10% -10%

✖The Propellant Mass Flow Rate refers to the amount of mass that flows through a given point in the rocket propulsion system per unit time.ⓘ Propellant Mass Flow Rate [ṁ]

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Formula

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Propellant Mass Flow Rate

Formula

`"ṁ" = ("A"_{"t"}*"P"_{"1"}*"γ")*sqrt((2/("γ"+1))^(("γ"+1)/("γ"-1)))/sqrt("γ"*"[R]"*"T"_{"1"})`

Example

`"11.32815kg/s"=("0.21m²"*"0.0037MPa"*"1.33")*sqrt((2/("1.33"+1))^(("1.33"+1)/("1.33"-1)))/sqrt("1.33"*"[R]"*"256K")`

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Propellant Mass Flow Rate Solution

STEP 0: Pre-Calculation Summary

Formula Used

Propellant Mass Flow Rate = (Nozzle Throat Area*Inlet Nozzle Pressure*Specific Heat Ratio)*sqrt((2/(Specific Heat Ratio+1))^((Specific Heat Ratio+1)/(Specific Heat Ratio-1)))/sqrt(Specific Heat Ratio*[R]*Temperature at Chamber)
ṁ = (A_t*P₁*γ)*sqrt((2/(γ+1))^((γ+1)/(γ-1)))/sqrt(γ*[R]*T₁)
This formula uses 1 Constants, 1 Functions, 5 Variables

Constants Used

[R] - Universal gas constant Value Taken As 8.31446261815324

Functions Used

sqrt - A square root function is a function that takes a non-negative number as an input and returns the square root of the given input number., sqrt(Number)

Variables Used

Propellant Mass Flow Rate - (Measured in Kilogram per Second) - The Propellant Mass Flow Rate refers to the amount of mass that flows through a given point in the rocket propulsion system per unit time.
Nozzle Throat Area - (Measured in Square Meter) - The nozzle throat area refers to the cross-sectional area of the narrowest part of a propulsion nozzle, known as the throat.
Inlet Nozzle Pressure - (Measured in Pascal) - Inlet Nozzle Pressure represents the pressure of the incoming air or propellant before it enters the combustion chamber or the turbine section.
Specific Heat Ratio - The specific heat ratio describes the ratio of specific heats of a gas at constant pressure to that at constant volume.
Temperature at Chamber - (Measured in Kelvin) - Temperature at Chamber typically refers to the temperature inside a closed chamber or enclosure.

STEP 1: Convert Input(s) to Base Unit

Nozzle Throat Area: 0.21 Square Meter --> 0.21 Square Meter No Conversion Required
Inlet Nozzle Pressure: 0.0037 Megapascal --> 3700 Pascal (Check conversion here)
Specific Heat Ratio: 1.33 --> No Conversion Required
Temperature at Chamber: 256 Kelvin --> 256 Kelvin No Conversion Required

STEP 2: Evaluate Formula

Substituting Input Values in Formula

ṁ = (A_t*P₁*γ)*sqrt((2/(γ+1))^((γ+1)/(γ-1)))/sqrt(γ*[R]*T₁) --> (0.21*3700*1.33)*sqrt((2/(1.33+1))^((1.33+1)/(1.33-1)))/sqrt(1.33*[R]*256)

Evaluating ... ...

ṁ = 11.328154115397

STEP 3: Convert Result to Output's Unit

11.328154115397 Kilogram per Second --> No Conversion Required

FINAL ANSWER

11.328154115397 ≈ 11.32815 Kilogram per Second <-- Propellant Mass Flow Rate

(Calculation completed in 00.020 seconds)

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< 4 Propellents Calculators

Propellant Mass Flow Rate

Go Propellant Mass Flow Rate = (Nozzle Throat Area*Inlet Nozzle Pressure*Specific Heat Ratio)*sqrt((2/(Specific Heat Ratio+1))^((Specific Heat Ratio+1)/(Specific Heat Ratio-1)))/sqrt(Specific Heat Ratio*[R]*Temperature at Chamber)

Oxidizer Mass Flow Rate

Go Oxidizer Mass Flow Rate = (Propellant Mixture Ratio*Propellant Mass Flow Rate)/(Propellant Mixture Ratio+1)

Fuel Mass Flow rate

Go Fuel Mass Flow Rate = Propellant Mass Flow Rate/(Propellant Mixture Ratio+1)

Propellant Mixture Ratio

Go Propellant Mixture Ratio = Oxidizer Mass Flow Rate/Fuel Mass Flow Rate

Propellant Mass Flow Rate Formula

What is Mass Flow Rate ?

Mass flow rate is the rate at which fuel and oxidizer are fed into a rocket's combustion chamber, measured in kilograms per second (kg/s). The higher the mass flow rate, the greater the thrust.

How to Calculate Propellant Mass Flow Rate?

Propellant Mass Flow Rate calculator uses Propellant Mass Flow Rate = (Nozzle Throat Area*Inlet Nozzle Pressure*Specific Heat Ratio)*sqrt((2/(Specific Heat Ratio+1))^((Specific Heat Ratio+1)/(Specific Heat Ratio-1)))/sqrt(Specific Heat Ratio*[R]*Temperature at Chamber) to calculate the Propellant Mass Flow Rate, The Propellant Mass Flow Rate is a crucial parameter that directly influences engine performance and thrust generation. It represents the rate at which mass (propellant) is expelled from the rocket engine, contributing to the generation of thrust. Propellant Mass Flow Rate is denoted by ṁ symbol.

How to calculate Propellant Mass Flow Rate using this online calculator? To use this online calculator for Propellant Mass Flow Rate, enter Nozzle Throat Area (A_t), Inlet Nozzle Pressure (P₁), Specific Heat Ratio (γ) & Temperature at Chamber (T₁) and hit the calculate button. Here is how the Propellant Mass Flow Rate calculation can be explained with given input values -> 11.32815 = (0.21*3700*1.33)*sqrt((2/(1.33+1))^((1.33+1)/(1.33-1)))/sqrt(1.33*[R]*256).

FAQ

What is Propellant Mass Flow Rate?

The Propellant Mass Flow Rate is a crucial parameter that directly influences engine performance and thrust generation. It represents the rate at which mass (propellant) is expelled from the rocket engine, contributing to the generation of thrust and is represented as ṁ = (A_t*P₁*γ)*sqrt((2/(γ+1))^((γ+1)/(γ-1)))/sqrt(γ*[R]*T₁) or Propellant Mass Flow Rate = (Nozzle Throat Area*Inlet Nozzle Pressure*Specific Heat Ratio)*sqrt((2/(Specific Heat Ratio+1))^((Specific Heat Ratio+1)/(Specific Heat Ratio-1)))/sqrt(Specific Heat Ratio*[R]*Temperature at Chamber). The nozzle throat area refers to the cross-sectional area of the narrowest part of a propulsion nozzle, known as the throat, Inlet Nozzle Pressure represents the pressure of the incoming air or propellant before it enters the combustion chamber or the turbine section, The specific heat ratio describes the ratio of specific heats of a gas at constant pressure to that at constant volume & Temperature at Chamber typically refers to the temperature inside a closed chamber or enclosure.

How to calculate Propellant Mass Flow Rate?

The Propellant Mass Flow Rate is a crucial parameter that directly influences engine performance and thrust generation. It represents the rate at which mass (propellant) is expelled from the rocket engine, contributing to the generation of thrust is calculated using Propellant Mass Flow Rate = (Nozzle Throat Area*Inlet Nozzle Pressure*Specific Heat Ratio)*sqrt((2/(Specific Heat Ratio+1))^((Specific Heat Ratio+1)/(Specific Heat Ratio-1)))/sqrt(Specific Heat Ratio*[R]*Temperature at Chamber). To calculate Propellant Mass Flow Rate, you need Nozzle Throat Area (A_t), Inlet Nozzle Pressure (P₁), Specific Heat Ratio (γ) & Temperature at Chamber (T₁). With our tool, you need to enter the respective value for Nozzle Throat Area, Inlet Nozzle Pressure, Specific Heat Ratio & Temperature at Chamber and hit the calculate button. You can also select the units (if any) for Input(s) and the Output as well.

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