Richard Nakka’s Experimental Rocketry Web Site


Solid Rocket Motor Theory – Two-phase Flow


Introduction to Two-phase Flow

Most solid rocket propellants produce combustion products that are a mixture of gases and condensed-phase particles (either liquid or solid) which is evident as visible smoke in the exhaust plume. Those propellants containing metals, such as aluminum or magnesium, generate oxides of the metals as condensed-phase combustion products. Metallic-compound oxidizers, such as potassium nitrate (KN) or potassium perchlorate (KP), generate condensed-phase products of particularly high molecular mass. As is seen by Equation 12 of Nozzle Theory web page, a higher molecular mass (M) of the products lowers the exhaust velocity and therefore overall performance. The KN-Sugar propellants produce a dense white cloud of potassium carbonate smoke. In fact, approximately 42% of the exhaust mass is condensed-phase matter !

The occurrence of solids or liquids in the exhaust leads to a reduction in performance for a number of additional reasons:

·    This portion of the combustion mass cannot perform any expansion work and therefore does not contribute to acceleration of the exhaust flow.

·    The higher effective molecular mass of these products lowers the Characteristic Velocity (c-star).

·    Due to thermal inertia, the heat of the condensed-phase is partly ejected out of the nozzle before transferring this heat to the surrounding gases, and is therefore not converted to kinetic energy. This is known as particle thermal lag.

·    Likewise due to the relatively large mass of the particles (compared to the gases), these cannot accelerate as rapidly as the surrounding gases, especially in that portion of the nozzle where flow acceleration is extremely high (throat region). Acceleration of the particles depends upon frictional drag in the gasflow, which necessitates a differential velocity between the particles and the gas. The net result is that the condensed-phase particles exit the nozzle at a lower velocity than the gases. This is referred to as particle velocity lag.

There are a number of factors that determine how significant two-phase flow losses have on a rocket motor’s performance. One important practical factor is nozzle contour, in particular at the throat region. A more gradual acceleration of the flow in the neighbourhood of the throat results in a reduction of thrust loss (Ref.1). Figure 1 illustrates the flow acceleration for the Kappa nozzle. The acceleration in the region of the throat (red dashed line) is extremely high, especially just aft, where it is maximum. Most of the particle lag, which is a strong function of acceleration, occurs in this region, thus the importance of designing a nozzle with a well-rounded contour at the throat, without any sharp changes in cross-section.

acceleration influence

Figure 1 -- Gas/particle acceleration for two-phase flow through Kappa nozzle

The size of the rocket motor as well as condensed-phase particle size both play an important role with regard to the influence of two-phase flow effects. This is illustrated in Figure 2, which plots the fraction of Characteristic Velocity loss with respect to:

·    Motor size (thrust)

·    Particle size

Note that the mass fraction of particles in the exhaust for this study was 0.25. For the sugar propellants, the mass fraction can be as high as 0.44.

Size influence

Figure 5 -- Influence of motor size and particle size on c-star
Excerpted from Ref.2

For example, for a 100 lb. (445 N.) thrust motor, the motor suffers a 6% loss in Characteristic Velocity if the average particle size is 1.5 micron, as shown by the orange dashed line.
It is clear from this plot that for amateur experimental motors, which are typically of 1000 lb. thrust or less, that two-phase flow losses can be significant, but are less significant for large "professional" motors.

In the following treatment of two-phase flow and how it affects the performance the various parameters related to rocket motor design, the assumption of zero particle lag is assumed. In other words, the particles are assumed to have the same temperature and same velocity as the surrounding gases as the products flow through the nozzle. This assumption serves to greatly simplify the two-phase problem in a practical way for rocket motor design. In reality, this is largely true for exhaust particles of one micron size (Ref.1). This is likely typical for sugar propellants, which produce a very fine smoke.


Rocket Motor Design Parameters for Two-Phase Flow

The following section provides the means for calculating thermochemical and performance parameters for a rocket motor that generates two-phase exhaust products. The superscript ʺ is applied to properties denoting that is is applicable to two-phase exhaust flow.

Effective Molecular Mass and Gas Constant

When the exhaust products consist of a mixture of gases and condensed-phase, the effective molecular mass, designed , is given by the number of gas moles, Ng, divided into the total mass of the exhaust products, mp:

The specific gas constant for a mixture of gases and condensed-phase, Rʺ, is given by the universal gas constant, , divided by the effective molecular mass of the products, :

The specific gas constant for the gas-only products, R, is related by the particle fraction, X, as shown in Eqn.2b:


Ratio of Specific Heats

The ratio of specific heats (also refered to as isentropic exponent) for a mixture of gases and condensed phase, kmix, is given by:

where Cp mix is the specific heat (at constant pressure) of the mixture of gases and condensed phase and is given by:




where m1/mp, m2/mp, etc are the mass fraction of each of the gaseous and condensed-phase products and Cp 1, Cp 2, etc. are the specific heats of gaseous products. Note that for condensed-phase products, the specific heat is denoted Cs. Specific heat is a function of temperature, so the values must be determined at the combustion temperature.

The specific heat for the two-phase mixture is related to that of the gas-only and condensed-phase products and by the particle fraction as shown by Eqn.4b:


The quantity kmix is valid for a static mixture of gases and condensed-phase products. What about the dynamic case where the products are flowing through the nozzle? A solution to this case is obtained from gasdynamics analysis, as given in References 3 & 4, with the derivation shown in the Appendix. In this case, the isentropic exponent for two-phase flow, , is given by:

where k is the gas-only products specific heat ratio and Ψ (Greek letter Psi) is the term X/(1-X), where X is the mass fraction of condensed-phase products. Note that Cs is for the mixture of condensed-phase products and Cp is for the mixture of gaseous products.

An alternate form of expression for kʺ is given by:

Eqn.3 and Eqn.5a can be shown (see Appendix) to be algebraically equivalent, such that = kmix.


Critical (throat) Velocity

At the nozzle throat, the velocity of the product flow is limited to mach one, or sonic velocity. With the assumption of two-phase mixture behaving as an ideal gas, the resulting equation is the same as for gas-only flow, except with the specific heat ratio and gas constant being those applicable to the two-phase mixture, and is given by Eqn.6a:

where T* is the critical (or throat) temperature of the flow. With the assumption of the two-phase mixture behaving as an ideal gas, the critical temperature is found using Eqn.4 of the Nozzle Theory web page, yielding Eqn.6b:

The critical velocity is of special interest to the rocket motor designer. As was shown by Eqn.2, the higher the fraction of condensed-phase, the lower the value of the gas constant, Rʺ. This reduces the value of v* which has the overall effect of lowering motor performance. The throat acts as a restrictor, or limiter, of gas velocity.


Characteristic Exhaust Velocity

In terms of the rocket performance parameters, the presence of condensed-phase products is reflected in a reduced Characteristic Exhaust Velocity (c-star or C*), due to the higher effective molecular mass of the gas/particle mixture. As was explained earlier, c-star is  figure of thermochemical merit for a particular propellant. With the assumption of two-phase mixture behaving as an ideal gas, the Characteristic Exhaust Velocity for a two-phase mixture is given by Eqn.7:


Exhaust Velocity

The exhaust velocity of a rocket nozzle, for two-phase flow, is given by Eqn.8:

where and are for two-phase products. Equation 8 is of identical form to that of gas-only flow, as seen by Eqn.12 of the Nozzle Theory web page.


Specific Impulse

The ideal specific impulse of a rocket motor is given by Eqn.9:

where c is the effective exhaust velocity and g is the acceleration of gravity ( sort of a conversion factor to give units of seconds). For a nozzle with optimum expansion ratio, the effective exhaust velocity is equivalent to the ideal exhaust velocity given by Eqn.8. As such, for two-phase flow, the ideal specific impulse is given by Eqn.10:


From Equation 2 of Thrust Theory web page, the thrust of a rocket motor is given by:


with the assumption of optimum nozzle expansion (Pe = Pa). The term ρ*A*v* represents the mass flow rate through the nozzle (* denotes critical or throat). Using Equation 7 of the Nozzle Theory web page, and based on the assumption that a two-phase mixture behaves as an ideal gas, the critical flow density at the nozzle throat is given by:


A* is the throat cross-sectional area. The critical velocity, v*, is given by Eqn.6 and exhaust velocity ve is given by Eqn.8.

From Eqn.6, Eqn.8, Eqn.11 and Eqn.12, the equation for thrust for two-phase flow is obtained (for optimum nozzle expansion) and is given by Eqn.13:

It is seen that the thrust equation for two-phase flow is of the same form as for gas-only flow given by Equation 3 of the Thrust Theory web page except having the two-phase parameters Rʺ and .


Thrust Coefficient

As explained in the Thrust Theory web page, the Thrust Coefficient determines the amplification of thrust due to gas expansion in the divergent portion of the nozzle as compared to the thrust that would be exerted if the chamber pressure acted over the throat area only. From equation 4 of the Thrust Theory web page, for a nozzle with optimum expansion:



The Thrust Coefficient for two-phase flow condition is therefore obtained from Eqn.13, and given by Eqn.14:





Chamber Pressure

The equation for Chamber Pressure for two-phase mixture is essentially the same as for the gas-only condition, as given by Equation 11 of the Chamber Pressure Theory web page. The only difference is with the values for specific heat ratio and gas constant, being those applicable for a two-phase mixture.



For a definition of the terms introduced in this equation, refer to the Chamber Pressure Theory web page.




Derivation of Eqn.5a (2-phase isentropic exponent)

Derivation of Eqn.5b (2-phase isentropic exponent)

Derivation of Eqn.8 (2-phase exhaust velocity)

Derivation of Eqn.13 (2-phase thrust)

Derivation of Eqn.15 (2-phase chamber pressure)

Show that Eqn.3 and Eqn.5 are algebraically equivalent


Example 1: For KNSU propellant, calculate the various two-phase thermochemical and performance parameters (ideal):

·       Molecular mass, specific gas constant and ratio of specific heats

·       Critical (throat) velocity, Characteristic Exhaust Velocity

·       Exhaust velocity and Specific Impulse (ideal expansion from 1000 psi chamber pressure)


Example 2: Utilizing KNSU propellant, calculate the ideal chamber pressure, thrust and thrust coefficient for the following motor conditions with optimum expansion:

·       Kn = 250

·       Throat diameter 12.0 mm (0.472 inch)



Reference Sources

1.     Gas Particle Flow in an Axisymmetric Nozzle, W.S.Bailey, E.N.Nilson, R.A.Serra and T.F.Zupnik, ARS Journal, June 1961

2.     Dynamics of Two-Phase Flow in Rocket Nozzles, M.Gilbert, J.Allport and R.Dunlap, ARS Journal, December 1962

3.     Mechanics and Thermodynamics of Propulsion, Philip G. Hill, Carl R.Peterson
Addison-Wesley Publishing Co.

4.     Recent Advances in Gas-Particle Nozzle Flows, Richard F.Hoglund, ARS Journal, May 1962

5.     Perturbation Analysis of One-dimensional Heterogeneous Flow in Rocket Nozzles, W.D.Rannie, Detonation and Two-Phase Flow edited by S.S.Penner & F.A.Williams, Academic Press, 1962

6.     Rocket Propulsion Elements, George P. Sutton, 4th Ed., John Wiley and Sons


Next -- Corrections for Actual Rocket Motors


Originally posted March 3, 2023

Last updated March 22, 2023

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