Numerical study of liquid propellants combustion for space applications

System design of liquid rocket engines must consider engine performance, weight, cost, and reliability requirements. A general design optimization framework has been developed in this paper to select the best system parameters for liquid rocket engines with gas-generator cycles. The object is to maximize the specific impulse and vacuum thrust-to-weight ratio of the engine with given system requirements and design assumptions by changing thrust-chamber pressure and mixture ratio. The system analysis, along with the engine weight estimation, is based on a modular scheme. Multidisciplinary design optimization formulations including multidisciplinary feasible and collaborative optimization are used, evaluated, and compared during the optimization process. Several techniques of multi-objective processing are also used to identify the Pareto frontier and the optimal compromise solutions. A proposed cryogenic-propellant engine using liquid oxygen and hydrogen with a gas-generator cycle is studied as a specific example. Moreover, uncertainties in the engine operation, such as thrust-chamber pressure and mixture ratio, are taken into account as random variables in the reliability-based optimization. Results are presented to illustrate the tradeoff between the engine performance and reliability requirements.

Optimization of System Parameters for Liquid Rocket Engines with Gas-Generator Cycles

A new engineering environment constructed for the purposes of analyzing and designing Reusable Launch Vehicles (RLVs) is presented. The new environment has been developed to allow NASA to perform independent analysis and design of emerging RLV architectures and technologies. The new Advanced Engineering Environment (AEE) is both collaborative and distributed. It facilitates integration of the analyses by both vehicle performance disciplines and life-cycle disciplines. Current performance disciplines supported include: weights and sizing, aerodynamics, trajectories, propulsion, structural loads, and CAD-based geometries. Current life-cycle disciplines supported include: DDT&E cost, production costs, operations costs, flight rates, safety and reliability, and system economics. Involving six NASA centers (ARC, LaRC, MSFC, KSC, GRC and JSC), AEE has been tailored to serve as a web-accessed agency-wide source for all of NASA's future launch vehicle systems engineering functions. Thus, it is configured to facilitate (a) data management, (b) automated tool/process integration and execution, and (c) data visualization and presentation. The core components of the integrated framework are a customized PTC Windchill product data management server, a set of RLV analysis and design tools integrated using Phoenix Integration's Model Center, and an XML-based data capture and transfer protocol. The AEE system has seen production use during the Initial Architecture and Technology Review for the NASA 2nd Generation RLV program, and it continues to undergo development and enhancements in support of its current main customer, the NASA Next Generation Launch Technology (NGLT) program.

/pdf/multi-disciplinary-analysis-for-future-launch-systems-using-2ygbt5fsto.pdf

Multi-Disciplinary Analysis for Future Launch Systems Using NASA's Advanced Engineering Environment (AEE)

Theaimofthisworkistoinvestigatenumericallythermalbarriercoating (TBC) systems applied to realistic rocket thrust chamber conditions. A global full parametric three-dimensional (3D) modeling approach for cooled rocket thrust chambers is presented to be able to simulate the §uidstructure interaction (FSI) phenomena involved. In a subsequent analysis step, realistic mechanical and thermal boundary conditions are extractedfromcriticaldesignregionsoftheglobalmodelandappliedtoa local¦niteelementmodel(FEM)toanalyzepossibleTBCdelaminations by means of a Fracture Mechanics (FM) approach.

/pdf/fluid-structure-interaction-analysis-applied-to-thermal-1nkqycnaov.pdf

Fluid-Structure Interaction Analysis Applied to Thermal Barrier Coated Cooled Rocket Thrust Chambers with Subsequent Local Investigation of Delamination Phenomena

A design method for a kerosene fuel-rich gas-generator of a liquid rocket engine using turbopumps to supply propellant was performed at a conceptual level. The gas-generator creates hot gases, enabling the turbine to operate the turbopumps. A chemical non-equilibrium analysis and a droplet vaporization model were used for the estimation of the burnt gas properties and characteristic chamber length. A premixed counter-flow flame analysis was performed for the prediction of the burnt gas properties, namely the temperature, the specific heat ratio and heat capacity, and the chemical reaction time. To predict the vaporization time, the Spalding model, using a single droplet in convective condition, was used. The minimum residence time in the chamber and the characteristic length were calculated by adding the reaction time and the vaporization time. Using the characteristic length, the design methods for the fuel-rich gas-generator were established. Finally, a parametric study was achieved for the effects of the O/F ratio, mass flow rate, chamber pressure, initial droplet temperature, initial droplet diameter and initial droplet velocity.

Conceptual Design for a Kerosene Fuel-rich Gas-generator of a Turbopump-fed Liquid Rocket Engine

The SCORES-II program is a conceptual-level engineering design and analysis tool for the performance prediction of liquid propellant rocket engines. The tool is written in the C++ programming language and can be compiled and executed on all major computing platforms (e.g. UNIX, PC, Mac). SCORES-II can be executed through a command line/console application or via the ModelCenter © environment with its filewrapper. The Analysis Server © filewrap allows for automated execution and integration with other disciplinary analysis tools. SCORES-II can support a wide variety of engine configurations. In addition to the built-in propellant options, the chemical equilibrium routine is capable of handling any generic fuel and oxidizer combination. Numerous options exist for sizing an engine, including thrust and throat area matching. An 'expert system' of engine efficiencies is also included, allowing for cycle, chemical reaction, injector/combustor, and nozzle influences on performance. The equations used for the performance calculations, equilibrium chemistry model, and engine sizing will be presented and discussed. Chemical equilibrium results obtained from implementing these equations will then be compared with results from another equilibrium code widely used in industry. Performance predictions from SCORES-II will be compared with known performance values for a number of existing engines. Sample results demonstrating SCORES-II's throttled engine analysis will be presented for a notional engine design. The tool's user-interface options will also be discussed. Finally, the paper will include a future work section detailing additional improvements and capabilities planned for future incarnations of SCORES.

John Bradford

Papers

SCORES-II Design Tool for Liquid Rocket Engine Analysis