Method and system for predicting rocket nozzle deformation during engine start-up and shut-down transients

Patent metadata

Field	Value
Publication number	US-9977848-B1
Application number	US-201414553234-A
Country	US
Kind code	B1
Filing date	Nov 25, 2014
Priority date	Jul 10, 2014
Publication date	May 22, 2018
Grant date	May 22, 2018

Abstract

Computational fluid dynamics (CFD) computations are performed at time increments using structural properties of the nozzle and flow properties of combustion products flowing through the nozzle. Each CFD computation accounts for movement of the wall geometry of the rocket nozzle due to the flowfield. Structural dynamics computations are performed at each time increment using the CFD computations in order to describe the movement of the wall geometry. Mesh dynamics computations at each time increment redefine the flowfield to account for the movement of the wall geometry. The mesh dynamics computations are based on a spring analogy process. The computations are iterated to solution convergence at each time increment with results being output to an output device.

First claim

What is claimed as new and desired to be secured by Letters Patent of the United States is:

1. A method of predicting deformation of a rocket nozzle due to rocket engine combustion products flowing through the rocket nozzle during times associated with engine start-up and engine shut-down, comprising the steps of:
providing structural parameters describing the rocket nozzle and wall geometry thereof;
providing flow properties generated by the rocket engine combustion products and entering the rocket nozzle during the times associated with engine start-up and engine shut-down;
providing a processing system that includes at least one processor for implementing a plurality of programmed processing steps and at least one output device for presenting results generated by said programmed processing steps;
said programmed processing steps including
performing computational fluid dynamics (CFD) computations at time increments using said structural parameters and said flow properties wherein said CFD computations describe a flowfield associated with the rocket engine combustion products flowing through the rocket nozzle, wherein each of said CFD computations accounts for movement of the wall geometry of the rocket nozzle due to said flowfield,
performing structural dynamics computations at each of said time increments using said CFD computations wherein said structural dynamics computations describe the movement of the wall geometry, and
performing mesh dynamics computations at each of said time increments to redefine said flowfield to account for the movement of the wall geometry, said mesh dynamics computations being based on a spring analogy process; and
outputting results of said programmed processing steps for at least one of said time increments using said at least one output device.

2. A method according to claim 1 , wherein said steps of performing are iterated for each of said time increments until a solution converges at each of said time increments.

3. A method according to claim 1 , wherein said flow properties include profiles of pressure of the rocket engine combustion products entering the rocket nozzle, profiles of temperature of the rocket engine combustion products entering the rocket nozzle, and profiles of chemical species concentration associated with the rocket engine combustion products entering the rocket nozzle.

4. A method according to claim 1 , wherein said CFD computations solve time-varying transport equations of continuity, species continuity equations, momentum equations, total enthalpy equations, turbulent kinetic energy equations, and turbulent kinetic energy dissipation equations.

5. A method according to claim 1 , wherein said step of outputting includes the step of displaying images of said results of said programmed processing steps.

6. A method according to claim 1 , wherein said spring analogy process includes a weighting function, said weighting function being greater than unity at the wall geometry of the rocket nozzle and transitioning to unity at a boundary layer edge in said flowfield.

7. A system for predicting deformation of a rocket nozzle due to rocket engine combustion products flowing through the rocket nozzle during times associated with engine start-up and engine shut-down, comprising:
a processing system that includes at least one processor for implementing a plurality of programmed processing steps and at least one database for storage of data needed by said programmed processing steps, wherein said data includes structural parameters describing the rocket nozzle and wall geometry thereof, and wherein said data includes flow properties generated by the rocket engine combustion products and entering the rocket nozzle during the times associated with engine start-up and engine shut-down,
said programmed processing steps including
performing computational fluid dynamics (CFD) computations at time increments using said structural parameters and said flow properties wherein said CFD computations describe a flowfield associated with the rocket engine combustion products flowing through the rocket nozzle, wherein each of said CFD computations accounts for movement of the wall geometry of the rocket nozzle due to said flowfield,
performing structural dynamics computations at each of said time increments using said CFD computations wherein said structural dynamics computations describe the movement of the wall geometry, and
performing mesh dynamics computations at each of said time increments to redefine said flowfield to account for the movement of the wall geometry, said mesh dynamics computations being based on a spring analogy; and
at least one output device coupled to said processor for presenting results generated by said programmed processing steps for at least one of said time increments.

8. A system as in claim 7 , wherein said steps of performing are iterated for each of said time increments until a solution converges at each of said time increments.

9. A system as in claim 7 , wherein said flow properties include profiles of pressure of the rocket engine combustion products entering the rocket nozzle, profiles of temperature of the rocket engine combustion products entering the rocket nozzle, and profiles of chemical species concentration associated with the rocket engine combustion products entering the rocket nozzle.

10. A system as in claim 7 , wherein said CFD computations solve time-varying transport equations of continuity, species continuity equations, momentum equations, total enthalpy equations, turbulent kinetic energy equations, and turbulent kinetic energy dissipation equations.

11. A system as in claim 7 , wherein said at least one output device is selected from the group consisting of devices that present said results in a viewable format and devices that store said results.

12. A system as in claim 7 , wherein said programmed process includes a weighting function, said weighting function being greater than unity at the wall geometry of the rocket nozzle and transitioning to unity at a boundary layer edge in said flowfield.

Assignees

• Nasa

Inventors

• Wang Ten-See

Classifications

• G06F30/23

  using finite element methods [FEM] or finite difference methods [FDM] · CPC title

• G06F30/15

  Vehicle, aircraft or watercraft design · CPC title

• F02K9/97 Primary

  Rocket nozzles (thrust or thrust vector control F02K9/80) · CPC title

• G06F17/10

  Complex mathematical operations {(function generation by table look-up G06F1/03; evaluation of elementary functions by calculation G06F7/544)} · CPC title

• G06F17/5009 Primary

  Physics · mapped topic