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edit:

Doh... nevermind, i found my mistake.. it should be

Tt = (1 + (gamma - 1) / 2 * M^2) * Tambient

-------------------------------------------------------

I'm trying to simulate an airbreathing engine and have encountered a strange problem:

When flying through the atmosphere I can calculate massflow through a given area by:

mdot = A * v * Rho

with mdot = massflow in kg/s, A = area in m^2, v = airspeed in m/2 and Rho = atmospheric density in kg/m^3

Now i did some testflights and compared the massflow calculated with the above formula to that calculated with another formula from this site: http://www.grc.nasa.gov/WWW/K-12/airplane/mflchk.html

which is

mdot = A * pt / sqrt(Tt) * (gamma / R) * M * (1 + (gamm - 1) / 2 * M^2) ^ (-(gamma+1)/(2*(gamma-1)))

where

Tt = total temperature calculated by

Tt = (1 + (gamma - 1) / gamma * M^2) * Tambient

(Tambient being the atmospheric temperature)

pt = total pressure calculated by

pt = pambient * (Tt / Tambient) ^ (gamma / (gamma - 1))

(pambient = atmospheric static pressure)

gamma = ratio of specific heat capacities as provided by the orbiter API (via oapiGetPlanetAtmConstants()), ~1.4 for earth

R = specific Gas constant also retrieved from oapiGetPlanetAtmConstants() ~286.9 for earth

M = freestream mach number

Strangely the two massflow results where significantly different with the second one starting a bit lower than the first at sealevel and slow airspeed and later getting significantly higher than the first one at higher altitudes and supersonic flight.

Would someone happen to have an idea whats wrong with my approach ?

Doh... nevermind, i found my mistake.. it should be

Tt = (1 + (gamma - 1) / 2 * M^2) * Tambient

-------------------------------------------------------

I'm trying to simulate an airbreathing engine and have encountered a strange problem:

When flying through the atmosphere I can calculate massflow through a given area by:

mdot = A * v * Rho

with mdot = massflow in kg/s, A = area in m^2, v = airspeed in m/2 and Rho = atmospheric density in kg/m^3

Now i did some testflights and compared the massflow calculated with the above formula to that calculated with another formula from this site: http://www.grc.nasa.gov/WWW/K-12/airplane/mflchk.html

which is

mdot = A * pt / sqrt(Tt) * (gamma / R) * M * (1 + (gamm - 1) / 2 * M^2) ^ (-(gamma+1)/(2*(gamma-1)))

where

Tt = total temperature calculated by

Tt = (1 + (gamma - 1) / gamma * M^2) * Tambient

(Tambient being the atmospheric temperature)

pt = total pressure calculated by

pt = pambient * (Tt / Tambient) ^ (gamma / (gamma - 1))

(pambient = atmospheric static pressure)

gamma = ratio of specific heat capacities as provided by the orbiter API (via oapiGetPlanetAtmConstants()), ~1.4 for earth

R = specific Gas constant also retrieved from oapiGetPlanetAtmConstants() ~286.9 for earth

M = freestream mach number

Strangely the two massflow results where significantly different with the second one starting a bit lower than the first at sealevel and slow airspeed and later getting significantly higher than the first one at higher altitudes and supersonic flight.

Would someone happen to have an idea whats wrong with my approach ?

Last edited: