wiki/ Motor Development Program
  1. News and Updates
  2. Program Justification
  3. Plan of Attack
  4. Current Issues List
  5. Design Choices
    1. Cooling Choice
    2. Fuel Choices
  6. Motor Operating Parameters
  7. Injector Element Design
  8. Engine Parameters
  9. Injector Plates
      1. Preliminary design parameters
      2. Preliminary Flow Calculations
    1. Topics for discussion
  10. Useful Mathematical Relationships and Information:
    1. Notation Guide
    2. Formulae
  11. References

News and Updates

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2012-02-14 - Digitalia Ordered

2012-02-08 - Motor Development Program and Pump Development Program Wiki Updates In Progress

2011-08-05 - Avionics Team Meeting



Program Justification

It is possible to insert a payload into orbit with big dumb boosters alone; the Japanese proved that with the Lambda 4S. Benjamin's google-fu has proved insufficient to determine much about the L4S' control system, other than that the upper stages used chlorofluorocarbon injection into the exhaust stream for thrust vectoring. Injecting liquids and gases into the motor exhaust seems an intuitively inefficient method to vector thrust (an intuition validated by the fact that rear-stage impacts on the forward stages drove the control system outside of its authority domains). India's Polar Satellite Launch Vehicle uses a flex nozzle for thrust vectoring on the third stage. The Polar Satellite Launch Vehicle also has an entire stage fueled by hydrazine and nitrogen tetraoxide (!).

It would probably be worthwhile to elide the requirements for a long-term vehicle development plan:

This requirement is driven by fuel considerations. If the orbital vehicle can hit well-optimized trajectories it will save buckets of fuel. This will reduce gravity losses and deliver either equivalent payloads with smaller rockets or larger payloads with equivalent rockets. - Hit some optimized point on the Isp/complexity tradeoff curve. - Reduce stage numbers to reduce complexity penalties.

Several scenarios are immediately apparent (and others are the result of fun with combinations):

It would be nice to have some objective criteria ("that's crazy" not being good engineering criteria) to evaluate solids vectoring vs. liquids vectoring. It seems obvious (again, intuitively that vectoring solids is a crazy plan; the unknown unknowns being a perceived larger set than the unknown unknowns of liquid development.

Plan of Attack

Current Issues List

Design Choices

Cooling Choice

Obvious Options: - Ablative - Regenerative - Open-cycle - Film

Fuel Choices

A exercise in combinatorics!

Motor Operating Parameters

A nice place to start designing the motor is a desired thrust level. Choosing too small of a thrust will make manufacturing difficult, and too large of a thrust will entail mass flows of lethal scales. There's an optimum in there somewhere, right?

Injector Element Design

It has been shown elsewhere (and examined to the satisfaction of Propulsion Team members) that a liquid flow resulting from the impingement of two jets directly opposed will flow axially down the chamber if the following holds:

\overset{\centerdot}{m}_f * v_f * cos(\gamma_f)= \overset{\centerdot}{m}_o * v_o * cos (\gamma_o)

where \gamma_f and gamma_o are the angles between the chamber axis and the fuel and oxygen flows respectively.

However, because we are working with a three-port injector element, the radial mass flows are not necessarily so simple. We can rewrite the above equation to account for the additional angle like so:

\overset{\centerdot}{m}_f * v_f * cos (\gamma_f) = 2 * \overset{\centerdot}{m}_o * v_o * cos(\gamma_z) * cos (\gamma_o)

A common injector parameter is the angle between the streams of oxygen and fuel flow in the plane of fuel flow. The value we have found and are designing from for this parameter is sixty degrees. We can describe the triangle formed by the fuel stream impingement angles and the angles at which the streams leave the injector plate like so:

180^o = 60^o + \gamma_f + \gamma_o

Choosing \gamma_f = 45^o arbitrarily, we can solve algebraically to show that\gamma_o = 75^o

\gamma_z can be derived from trigonometric and algebraic manipulations on the 3 port injector element maths above. The derivation is left as an exercise to the reader, but the final result is as shown:

\gamma_z = cos^{-1} \left[\frac{\overset{\centerdot}{m_f}v_fcos(\gamma_f)}{2\overset{\centerdot}{m_o}v_ocos(\gamma_o)} \right]

There are not-immediately-obvious singularities to avoid, but the values of gammaf and gammao set equal to 60o work well.

Engine Parameters

Injector Plates

Some random note on this:

Pressure drop across injectors - Typically 15-25% of chamber pressure. High pressure drops increase stability (RPE p.284). Seems logical. Some sources cite delta p across injector ~41% of chamber pressure to further reduce feed system induced combustion instability.

Angle of injection - Set to get axial flow after impingement, based on mass flow and angle. i.e. "Resultant momentum at the point of impingement between the fuel and oxidizer flow is axially directed"

What is a good injection velocity? One publication cited a study of 5 m/s up to 50 m/s, over a range of chamber pressures, vs. stability. Less than 18-20 m/s cited stability issues. This is likely due to fuel entering the chamber at speeds less than that of the flame front.


Preliminary design parameters

delta P (injector) = 41% of Pc

Injection velocity = 30 m/s

Injection type: OFO

Impingement angle = Fuel: 45 deg.; Oxygen: 11.8586 deg.

Pre-Impingement distance = 5 mm

Orifice L/D ration = 18.4 or 80 (decide this)

Preliminary Flow Calculations

O/F ratio =~ 2.35

RPE (277):

Q = A * \sqrt{\frac{2\Delta p}{\rho}}

Rearranging allows us to solve for total hole area in injector plate satisfying desired fuel flow and oxygen flow:

Ao = 1.158 E-6 m2

Af = 6.123 E-7 m2

The new script version computes these values and more, including from an input of number of impingement elements, the individual hole sizes.

Topics for discussion

Useful Mathematical Relationships and Information:

Notation Guide

Symbol Meaning
R, R' Specific Gas Constant, Gas Constant.
\overset{\centerdot}{w} Mass flow (combined fuel and oxidizer).
A_i Cross-sectional area at a point in the engine.
v_i Gas velocity at a point in the engine.
V_i Gas specific volume.
P_i Gas pressure.
T_i Temperature.
_c Subscript denoting engine chamber.
_t ... engine throat.
_e ... exhaust exit.
_a ... ambient pressure.
g Gravitational acceleration.
k Ratio of specific heats at constant pressure and volume. Thermodynamic constant for specific gases.
{\eta} Thermal efficiency of the motor. Function of pressure and temperature ratios.
N_m Mach number (dimensionless ratio of speed to local speed of pressure wave propagation).


R = \frac{R'}{M} whereR' is the universal gas constant andM the average molecular weight of the exhaust gases which can be found here:

\overset{\centerdot}{w} = \frac{A_tv_t}{V_t} (RPE 3-24)

\overset{\centerdot}{w} = \frac{Fg}{c} (RPE p. 52). This equation gives optimum fuel consumption as a function of thrust and exhaust speed c.

v_t = \sqrt{gkRT_t} = \sqrt{\frac{2gkRT_c}{k+1}} (RPE 3-23)

v_e = \sqrt{\frac{2*k}{k-1}*\frac{R'T_c}{M}* \left(1-\frac{P_e}{P_c} \right)^\frac{k-1}{k}}

N_m^2 = \left(\frac{2}{k-1} \right) \left[\left(\frac{P_c}{P_e}\right)^\frac{k-1}{k}-1 \right] ( 1-29)

T_t = T_c \left(\frac{2}{k+1} \right) (RPE 3-22)

V_t = V_c \left(\frac{k+1}{2}\right)^\frac{1}{k-1} (RPE 3-21)

P_t = P_c * \left (1+\frac{k-1}{2} \right) ^\frac{-k}{k-1} (Isentropic compression? Sourced from

V_c = \frac{RT_c}{P_c} (Ideal Gas Law)

V_e = V_c * \left( \frac{P_c}{P_e} \right) ^\frac{1}{k} (RPE 3-6 and p. 52)

A_i = \frac{\overset{\centerdot}{w}V_i}{v_i} (RPE 3-24)

A_t = \frac{\overset{\centerdot}{w}}{P_t}*\sqrt{\frac{R'T_t}{Mk}} (Braeunig 1.26)

A_e = \frac{A_t}{N_m} * \left[\frac{1+\frac{k-1}{2}*N_m^2}{\frac{k+1}{2}} \right]^\frac{k+1}{2*(k-1)} (Braeunig 1.30)


Liquid Rocket Engine Combustion Instability by Vigor Yang, Yang, William E. Andersen (Editor)

Rocket Propulsion Elements by George P. Sutton, Oscar Biblarz

Spray Characteristics of Impinging Jet Injectors at High Back-Pressure