Jet Propulsion/Fluid Mechanics: Wikis

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Up to date as of January 23, 2010
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< Jet Propulsion

Contents

Units & symbols

Primary units

Table 1.1: Primary Units
Quantity Name Symbol
Length meter m
Mass kilogram kg
Time second s
Temperature Kelvin K
Substance quantity mole mol

Derived units

Table 1.2: Derived Units
Quantity Name Symbol In fundamental units
Frequency Hertz Hz s-1
Force Newton N m • kg s-2
Pressure, stress Pascal Pa N • m-2
Energy, work, heat Joule J N • m
Power Watt W J• s-1
Heat capacity, entropy Joule per kelvin J K-1
Specific heat capacity,specific entropy Joule per kilogram kelvin J kg-1 K-1
Specific energy Joule per kilogram J kg-1
Thermal conductivity Watt per meter kelvin W m-1 K-1

Standard Atmosphere

Table 1.3: Standard Atmosphere in SI Units
Attribute Symbol Value
Pressure, Sea level P0 101,325 Pa
Temperature T0 288.15K, 15oC
Acceleration due to gravity g0 9.80665 m s-2
Air density ρ0 1.225 kg m-3
Kinematic viscosity ν0 1.46070x10-5 m2 s-1
Absolute viscosity μ0 1.7894x10-5 m2 s-1
Temperature lapse rate, (sea level to isothermal, 0-11km) -6.5 K km-1
Gas constant R 287.074 J kg-1 K-1
Specific heat constant volume cV 717.986J kg-1 K-1
Specific heat, constant pressure cP 1004.76 J kg-1 K-1
Specific heat ratio γ 1.4
Speed of sound, Sea level C0 (= 20.05*sqrt(T)) 340.3065 m s-1

Symbols

Table 1.5: Standard Symbols
Symbol Definition
M Mach number
M* Velocity/acoustic state where M=1.0
P Pressure
q Dynamic pressure
R Gas constant
T Temperature
W Mass flow rate
WTAP Flow parameter
γ Specific heat ratio
ρ Density
sub/superscripts
0 ground level
s static/stream
t total (isentropic stagnation)
x in front of normal shock
y behind normal shock
* at M=1

Gas Laws

Internal energy u = cv T
Enthalpy h = cp T
Kelvin to celsius oK = oC+273.15
Perfect Gas: P V = m R T
At constant temperature: P1/ P2 = V1/ V2
At constant pressure V1/ V2 = T1/ T2
At constant volume P1/ P2 = T1/ T2
Reversible adiabatic process P1V1γ=P2V2 γ
P1/ P2 = ( V2/ V1 )γ
T1/ T2 = ( V2/ V1 )γ-1
T2/ T2 = ( P2/ P1 )(γ-1)/γ
P1/ P2 = ( ρ1/ ρ2 )γ
Polytropic process: P1V1n=P2V2 n
P1/ P2 = ( V2/ V1 )n
T1/ T2 = ( V1/ V2 )1-n
T1/ T2 = ( P1/ P2 )(n-1)/n;
Bernoulli equation P/ρ + V2/2 + Z = constant
Steady flow equation q + h + V2/2 + Z = constant
Velocity of sound in perfect gas C = \sqrt{\gamma R T}
Specific heats R = cP - cV
γ = cP / cV
Mach Number M = \frac{V}{C}
Compressible flow
M^2 = \frac{2}{\gamma-1} \frac{u^2}{c_P T}
 \frac{P_t}{P_s} = \left[1+\frac{\gamma -1 }{2}M^2 \right]^\frac{\gamma}{\gamma-1}
 \frac{T_t}{T_s} = \left[1+\frac{\gamma -1 }{2}M^2 \right]
 \frac{\rho_t}{\rho_s} = \left[1+\frac{\gamma -1 }{2}M^2 \right]^\frac{1}{\gamma-1}
 \frac{A}{A^*} = \frac{1}{M}\left[\frac{1+\frac{\gamma -1 }{2}M^2}{\frac{\gamma +1 }{2}} \right]^\frac{\gamma+1}{2(\gamma-1)}
 M^* = M \sqrt{\frac{\frac{\gamma+1 }{2}}{1+\frac{\gamma-1 }{2}M^2}}
\theta = \frac{T}{T_0}

Adiabatic Processes

An adiabatic process is a thermodynamic process in which no heat is transferred to or from the working fluid. In an ideal gas turbine (Brayton cycle) the compression and expansion processes are adiabatic. We define θ as the pressure ratio in a process relating to the ambient conditions:

 \theta_2 = \left(\frac{P_2}{P_0}\right)


Then for adiabatic compression the temperature ration τ is:

 \tau_2 = \frac{T_2}{T_0} = \theta_2^\frac{\gamma-1}{\gamma}


For air γ = 1.4 so (γ-1)/ γ= 0.286

A log-log plot simplifies the analysis for quick engineering calculations.

Figure 1.3: Adiabatic process, air γ=1.4, on log-log plot
Example 1.2: Adiabatic and isobaric processes
Air at standard sea level conditions is compressed to 30 bar adiabatically. (0-3); heated to 1700K at constant pressure (3-4) and then expanded back to 1 bar adiabatically (4-5). What's the final temperature and how much heat is added in process 3-4?
See figure 1.3

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Kinetic energy to fluid temperature

Changing the velocity of the fluid by pressure changes simultaneously changes the temperature. Compression work raises the apparent temperature of the fluid. We can relate Mach number to the internal energy of the fluid:

Aerodynamic Analysis

Aerodynamic analysis attempts to progressively analyze the flows in the aerodynamic stages. Practical design includes substantial theoretical, computational and experimental analysis.

The stagnation or total temperatures and pressures are needed to measure the energy additions in high speed gas flows that occur in gas turbines. Using the Mach number allows us to factor in the compressibility of the gas.

Compressible fluid flow equations

See [NACA 1135: Equations, Tables, and Charts for Compressible Flow].

Continuity

In a steady flow, for any two sections of the flow on a stream tube

ρ1u1A1 = ρ2u2A2

or in differential form

 \frac{d}{dx} (\rho u A) = 0

or

 \frac{d\rho}{\rho} + \frac{du}{u} + \frac{dA}{A} = 0

Where ρ is the density of the fluid

u is the internal energy per unit mass of the fluid

and

A is the cross-sectional Area for the tube or channel

Momentum

The net force on the control volume matches the momentum change in the fluid

dp=-ρ u du

Energy

The change in enthalpy is balanced by the change in kinetic energy

dh + u du=0

Where h is enthalpy per unit mass, u + pv, pv being the product of pressure and volume

Enthalpy

The enthalpy of a gas h at temperature T is

h = cpT

where cp is the constant pressure specific heat of the gas. For air cp is about 1.005 kJ/kg K.

Entropy

Entropy change in a process can be expressed as

 s_2-s_1 = c_v \ln \left(\frac{T_2} {T_1}\right) - R \ln \left(\frac{\rho_2} {\rho_1}\right)

or in more conveniently terms of pressures

 s_2-s_1 = c_p \ln \left(\frac{T_2} {T_1}\right) - R \ln \left(\frac{P_2} {P_1}\right)

Stagnation temperature

Stagnation temperature is the temperature of the gas if it is brought to rest adiabatically. Adding the kinetic energy to the internal energy of the gas we get the relation

h_t = c_p T_t = c_p T + \frac{u^2}{2}
T_t = T + \frac{u^2}{2 c_p}

where Tt is the total(stagnation temperature of the flow.

The total enthalpy relation encapsulates the energy changes in isentropic compressors and turbines. To add energy to the flow the gas is put through a relative deceleration process against the compressor and diffuser surfaces and gains energy. To extract energy the gas is accelerated against nozzles and turbine buckets.

The Mach number of the flow is

 M = \frac{u}{a}=\frac {u}{\sqrt{\gamma R T}}

Substituting

T_t = T + \frac {\gamma R T M^2 }{ 2 c_p}


since R= cp - cv and γ = cp / cv

T_t = T\left(1 + \frac {\gamma M^2 }{2} (c_p-c_v)/c_p\right)
\frac{T_t}{T}=\left(1 + \frac {\gamma-1}{2}M^2\right)

Where γ (greek letter gamma) is the adiabatic expansion coefficent between pressure and volume p_1V_1^\gamma = p_2V_2^\gamma

This is the temperature if the gas is brought to rest adiabatically.

Isentropic stagnation pressure

\frac{p_t}{p}=\left(\frac{T_t}{T}\right)^{\gamma/(\gamma-1)}
\frac{p_t}{p}=\left(1 + \frac {\gamma-1}{2}M^2\right)^{\gamma/(\gamma-1)}

Isentropic stagnation density

\frac{\rho_t}{\rho}=\left(\frac{T_t}{T}\right)^{1/(\gamma-1)}
\frac{\rho_t}{\rho}=\left(1 + \frac {\gamma-1}{2}M^2\right)^{1/(\gamma-1)}

Duct flow

A steady inviscid adiabatic quasi-one dimensional flow obeys the following equations:

Differential continuity equation

d (ρ u A) =0

Differential momentum equation

dp=-ρ u du

Differential energy equation

dh + u du=0

Rearranging continuity

\frac{d\rho}{\rho}+\frac{d u}{u}+\frac{d A}{A}=0

Rewrite momentum equation

\frac{d p}{\rho}=\frac{dp}{ d \rho }\frac{d \rho}{\rho} = -u du
\frac{dp}{ d \rho } \equiv \frac{\part p}{\part \rho}

The velocity of sound is:

a =(dp / dρ)1/2

Rearranging and substituting:

a2=(dp / dρ)
a2 dρ / ρ = -u du
\frac{d \rho}{\rho}= -\frac{u du}{ a^2 }= -\frac{u^2}{ a^2 } \frac{du}{u} = -M^2 \frac{du}{u}

Substituting into continuity equation

M^2 \frac{du}{u}-\frac{d u}{u}-\frac{d A}{A}=0

We get the area velocity equation:

\frac{d A}{A} =(M^2-1) \frac{du}{u}

Thus for acceleration (positive du/u) the area must decrease for Mach numbers below 1 and increase for Mach numbers above 1.

The relationship between Mach number and duct area related to the throat area A* is:

\frac{A}{A^*}= \frac {1}{M} \left[ \frac{2}{\gamma+1}\left(1+\frac{\gamma-1}{2}M^2\right) \right]^{\frac{\gamma+1}{2(\gamma-1)}}

The temperature relation is

\frac{T}{T_t}= \left[ \frac{2}{\gamma+1}\left(1+\frac{\gamma-1}{2}M^2\right) \right]^{-1}

the pressure relation

\frac{p}{p_t}= \left[ \frac{2}{\gamma+1}\left(1+\frac{\gamma-1}{2}M^2\right) \right]^{-\frac{\gamma}{2(\gamma-1)}}

and the density relation

\frac{\rho}{\rho_0}= \left[ \frac{2}{\gamma+1}\left(1+\frac{\gamma-1}{2}M^2\right) \right]^{-\frac{1}{2(\gamma-1)}}

The figure below shows these relationships for air with γ of 1.4.

Isentropic flow through a duct

A fully expanded gas would approach a Mach number of infinity as it's temperature drops to absolute zero.

Mass flow

The figure above shows this exchange for a fluid with γ=1.4 undergoing an adiabatic expansion. Sonic velocity (Mach 1) is achieved when the pressure drops to 0.528 and the area for a particular mass flow is minimum at this Mach number. The flow at this condition is said to be choked and any further reductions in duct area will not produce acceleration of the stream. The mass flow per unit area is

 \dot{m} = A \rho v
 \dot{m} = A \rho_0 \left[ \frac{2}{\gamma+1}\left(1+\frac{\gamma-1}{2}M^2\right) \right]^{-\frac{1}{\gamma-1}} V

Nozzles

A nozzle converts internal energy of the gas into directed kinetic energy by expanding along a pressure gradient.

As the gas expands initially the volume increment is smaller than the velocity increment and the stream tube converges. A M=1 the effects balance and for M>1 the differential volume increase is greater than the velocity increase and a divergent stream is needed. The narrowest section of the nozzle is called the "throat".

Decreasing the pressure at the exit of a nozzle of fixed geometry increases the exit velocity until the the velocity in the smallest section of the nozzle becomes sonic. The nozzle is then said to be "choked" and further reduction of the exit pressure has no effect on the flow upstream of the throat.

The maximum exit velocity depends on the energy content of the source gas.

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Choked flow

Choked flow is the maximum flow that can pass through a passage for a given initial total conditions. Boundary layer effects further limit the flow through real nozzles.

Diffusers

A diffuser converts relative kinetic energy into pressure.

An ideal diffuser would recover the stagnation pressure, but practical diffusers cannot bring the fluid velocity to zero and have losses. The pressure recovered by such a diffuser is:

πd = pt2 / pt0

A subsonic diffuser is a divergent passage. Diffusers operate in an adverse pressure gradient regime and the boundary layer development must be carefully managed to avoid flow separation. Boundary layers can be energized by extraction or aspiration but this has energy and complexity costs.

Achieving stable supersonic diffusion without shockwaves is almost impossible, since instabilities become rapidly magnified as the flow can rapidly snap to become subsonic via a normal shockwave and accelerate in the convergent passage. Usually multiple inclined shockwaves are employed to minimize entropy rise.

Shocks

The shock is a thin boundary across which heat transfer and viscous heating make the flow subsonic. The isentropic relations above are not applicable across a shock wave. The total temperature across a shock (normal to the shock surface) remains constant but the total pressure is lost. The loss depends on the incident Mach number.

Normal shock

The Mach number M2 after the shock is:

M_2^2=\frac{1+[(\gamma-1)/2]M_1^2}{\gamma M_1^2-(\gamma-1)/2}

A higher incident Mach number will transition to a smaller downstream subsonic Mach number.

The density & velocity relation

\frac{\rho_2}{\rho_1}=\frac{u_1}{u_2}=\frac{(\gamma+1)M_1^2}{2+(\gamma-1)M_1^2}

the pressure relation

\frac{p_2}{p_1}=1+\frac{2 \gamma }{\gamma+1}(M_1^2-1)

and the temperature relation

\frac{T_2}{T_1}=\frac{h_2}{h_1}=\left[1+\frac{2 \gamma }{\gamma+1}(M_1^2-1)\right] \frac{2+(\gamma-1)M_1^2}{(\gamma+1)M_1^2}

Inclined shock


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