Lecture P5: Rocket Performance


General comments

I covered part of Chapter V of the notes and Chapter VI. This included 1) Principal figures of merit for rocket performance (thrust, specific impulse and weight), 2) discussion of the general layout of rocket motor components, and 3) Given typical design parameters (nozzle throat and exit areas, and combustor chamber conditions) how to calculate the principal figures of merit. We did onr PRS questions (PRS #1)

Next lecture we will cover the last part of rockets: given the principal figures of merit, how to calculate overall vehicle performance. Please review Chapter V of the notes.


Responses to 'Muddiest Part of the Lecture Cards'

(14 respondents)

1) Why for Isp=ue/g is g = 9.8 always (even on other planets)? (2 students) This is just the way Isp is defined. It is the thrust one would get for a unit weight flow rate of propellant where the weight is as measured here on earth.

2) How do you deal with varying combustion temps in the combustion chamber of the rocket engine in terms of calculations? (1 student) Do you mean varying as a function of location or time? In terms of location, we typically assume that the gases are well mixed (i.e. all at the same temperature). In terms of any variation with time, this would require solving the rocket equation for small increments of time, where the thrust was changed for each increment.

3) Would a pressure of 300atm inside the combustion chamber not make the Ae(po-pe) term in the thrust significant even when compared to mdot * ue? (1 student) Not, not typically. Consider the space shuttle main engine: A/A* = 77.5, pc=20.5MPa. Assuming a gamma of 1.4, then Mexit = 6.6 and pc/pe = 2855 based on isentropic flow relations. So the exit pressure would be 0.007MPa or about 7% of standard sea-level atmospheric pressure. The message: the flows are greatly expanded in rocket nozzles. (And even though the temperature starts out at 3000K, the isentropic temperature ratio is 9.7 -- so the static temperature is around room temperature. Of course the stagnation temperature is still 3000K.)

4) How do types of fuels and engines affect Isp? (1 student) In terms of fuels we haven't yet shown how to connect Tc and Pc to Isp (but we will next lecture). But once this connection is made, then what is required is a calculation or measurement of the temperature and pressure changes when various reactants are brought together and combusted. This is something we won't get to in class. It requires thermodynamics, fluid dynamics and chemistry. Further in terms of different types of engines (e.g. liquid propellant, solid propellant, ion engines, nuclear engines, etc.) we will only discuss liquid propellant engines. More details can be found in Hill and Peterson Mechanics and Thermodynamics of Propulsion (available in the library).

5) Pegasus launch? (1 student)You can check it out here: http://www.orbital.com/SpaceLaunch/Pegasus/index.html Note that David W. Thompson, the Chairman and Chief Executive Officer, and Orbital Co-Founder is a graduate from our department. And Professor Kerrebrock sits on the Board of Directors.

6) I will need to do a review of thermo from last term. (1 student) Yes you will.

7) So max endurance gives us maximum power required and maximum ranges gives us what? (1 student) You are a little confused. Maximum endurance is at the minimum-power-required point. Maximum range is at the minimum drag point.

8) No mud (6 students).