## Richard Nakka's Experimental Rocketry Web Site ### Solid Rocket Motor Theory -- Nozzle Theory

#### Nozzle Theory

The rocket nozzle can surely be described as the epitome of elegant simplicity. The primary function of a nozzle is to channel and accelerate the combustion products produced by the burning propellant in such as way as to maximize the velocity of the exhaust at the exit, to supersonic velocity. The familiar rocket nozzle, also known as a convergent-divergent, or deLaval nozzle, accomplishes this remarkable feat by simple geometry. In other words, it does this by varying the cross-sectional area (or diameter) in an exacting form.
The analysis of a rocket nozzle involves the concept of "steady, one-dimensional compressible fluid flow of an ideal gas". Briefly, this means that:

• The flow of the fluid (exhaust gases + condensed particles) is constant and does not change over time during the burn

• One-dimensional flow means that the direction of the flow is along a straight line. For a nozzle, the flow is assumed to be along the axis of symmetry
. • The flow is compressible. The concept of compressible fluid flow is usually employed for gases moving at high (usually supersonic) velocity, unlike the concept of incompressible flow, which is used for liquids and gases moving at a speeds well below sonic velocity. A compressible fluid exhibits significant changes in density, an incompressible fluid does not.

• The concept of an ideal gas is a simplifying assumption, one that allows use of a direct relationship between pressure, density and temperature, which are properties that are particularly important in analyzing flow through a nozzle.

Fluid properties, such as velocity, density, pressure and temperature, in compressible fluid flow, are affected by

1. cross-sectional area change
2. friction
3. heat loss to the surroundings
The goal of rocket nozzle design is to accelerate the combustion products to as high an exit velocity as possible. This is achieved by designing the necessary nozzle geometric profile with the condition that isentropic flow is to be aimed for. Isentropic flow is considered to be flow that is dependant only upon cross-sectional area -- which necessitates frictionless and adiabatic (no heat loss) flow. Therefore, in the actual nozzle, it is necessary to minimize frictional effects, flow disturbances and conditions that can lead to shock losses. In addition, heat transfer losses are to be minimized. In this way, the properties of the flow are near isentropic, and are simply affected only by the changing cross-sectional area as the fluid moves through the nozzle.

Typical nozzle cross-sectional areas of particular interest are shown in the figure below The analysis of compressible fluid flow involves four equations of particular interest:

1. Energy
2. Continuity
3. Momentum
4. The equation of state

The energy equation is a statement of the principle of conservation of energy. For adiabatic flow between any two points, x1 and x2 , it is given by  where h represents enthalpy of the fluid (which can be considered the energy available for heat transfer), v is the flow velocity in the x-direction, Cp is the effective heat capacity of the fluid, and T is the fluid temperature.
This equation provides valuable insight into how a rocket nozzle works. Looking at the first two terms shows that the change (decrease) in enthalpy is equal to the change (increase) in kinetic energy. In other words, heat of the fluid is being used to accelerate the flow to a greater velocity. The third term represents the resulting change (decrease) in temperature of the flow. The heat capacity may be approximated to be constant, and is a property determined by the composition of the combustion products.

It is apparent, then, that the properties of a fluid (e.g. temperature) are a function of the flow velocity. In describing the state of a fluid at any point along its flow, it is convenient to consider the stagnation state as a reference state. The stagnation properties may be considered as the properties that would result if the fluid were (isentropically) decelerated to zero velocity (i.e. stagnant flow).

The stagnation temperature, To, is found from the energy equation (by setting v2=0) to be equation 1

For an isentropic flow process, the following important relationship between stagnation properties for Temperature, Pressure, and Fluid Density hold equation 2

where k is the all-important ratio of specific heats, also referred to as the isentropic exponent, defined as Both Cp and R (specific gas constant) are properties determined by the composition of the combustion products, where R = R'/ M, where R' is the universal gas constant, and M is the effective molecular weight of the combustion products. If the combustion products contain an appreciable percentage of condensed phase particles (smoke), the value of the effective molecular weight, M, must account for this. As well, the proper k must be used which takes into account two-phase flow. The determination of k and M for the combustion products is detailed in the Technical Notepad #1 Web Page.

The local sonic velocity, a, and the Mach number, M, (defined as the ratio of the flow velocity to the local sonic velocity), is given by equation 3

From equations 1,2 & 3, the relationship between the stagnation temperature (also referred to as total temperature) and Mach number may be written as equation 4

It can be shown from the first and second laws of thermodynamics, for any isentropic process, that equation 5

From equations 4 & 5, and from the equation of state for an ideal gas, , the relationship between stagnation pressure; density and Mach number may be expressed as given in the following two equations equation 6 equation 7

Equations 4, 6 & 7 are particularly useful, as these allow each property to be determined in a flow if the Mach number and the stagnation properties are known. The stagnation (or total) properties To, Po, and o are simply the properties that are present in the combustion chamber of the rocket, since the flow velocity is (considered to be) zero at this location. In other words, To is the combustion temperature of the propellant (AFT), Po is the chamber pressure, and o is the density of the combustion products under chamber conditions.

Another important stagnation property is the stagnation enthalpy. This is obtained from the energy equation (by setting v2=0) equation 8

Physically, the stagnation enthalpy is the enthalpy that would be reached if the flow (at some point) were somehow decelerated to zero velocity. It is useful to note that the stagnation enthalpy is constant throughout the flow in the nozzle. This is also true of the other stagnation properties (temperature, pressure, and density).

The second of the four equations of interest regarding compressible fluid flow, as discussed earlier, is the continuity (or conservation of mass) equation, which is given by equation 9

where A is the nozzle cross-sectional area, v is the velocity of the flow. This equation simply states that the mass flowing through the nozzle must be constant. The "star" (asterisk) signifies a so-called critical condition, where Mach number is unity, M=1 (flow velocity is equal to the speed of sound). The importance of the critical condition will soon be made apparent.

Taking equations 3, 4, 7 & 9, it is possible to express the area ratio, A/A*, in terms of the Mach number of the flow. The area ratio is simply the cross-sectional area at any point (x) in the nozzle, to the cross-sectional area where the critical condition exists (M=1) equation 10

From equations 8 & 9, the flow velocity at the nozzle exit can be expressed as equation 11

where subscripts e and x signify exit and any point x along the nozzle axis, respectively. This equation can then be put into the far more useful form with the aid of the energy equation and the definition of k, as well as equation 2. equation 12

This equation is one of the most useful, as it allows the nozzle exit velocity to be calculated. In summarizing, it is necessary to know

A better understanding of the nozzle behaviour may be obtained by looking closely that this equation. It may be seen that

The ratio between the throat area, A*, and any downstream area in the nozzle, Ax, at which pressure Px prevails can be conveniently expressed as a function of the pressure ratio, Px /Po, and k. By noting that at the throat M is unity, and using equations 2, 3, 4, 7 & 12, leads to equation 13 equation 14

This is known as the nozzle design condition. For such a condition maximum thrust is achieved (derivation). For this design, the area ratio Ae /A* is known as the all-important Optimum Expansion Ratio.

For a highly informative explanation on convergent-divergent nozzle operation, in particular choked flow and shock formation, visit the Nozzle Applet website (includes a simulation).

#### Worked Examples Example #1 - Calculate nozzle area ratio (A/A*) with varying Mach number and plot on a graph
Worked Example #1

Example #2 - Calculate flow properties Temperature, Pressure and Density with varying Mach number and plot on a graph
Worked Example #2

Example #3 - Calculate nozzle Optimum Expansion ratio for a rocket motor operating at 65 atmospheres chamber pressure expanding to ambient air.
Worked Example #3

Example #4 - Calculate nozzle flow exit velocity for a rocket motor operating at 68 atmospheres chamber pressure, expanding to ambient air, using KNSB as propellant.
Worked Example #4

Next -- Rocket Motor Thrust 