Flow swirl impact at the rocket engine nozzle inlet on the flow coefficient

The recent growth of private options in launch vehicles has substantially raised price competition in the space launch market. This has increased the need to deliver reliable launch vehicles at reduced engine development cost, and has led to increased industrial interest in reduced order models. Large-scale liquid rocket engines require high-speed turbopumps to inject cryogenic propellants into the combustion chamber. These pumps can experience cavitation instabilities even when operating near design conditions. Of particular concern is rotating cavitation, which is characterized by an asymmetric cavity rotating at the pump inlet, which can cause severe vibration, breaking of the pump and loss of the mission. Despite much work in the field, there are limited guidelines to avoid rotating cavitation during design and its occurrence is often assessed through costly experimental testing. This paper presents a source term based model for stability assessment of rocket engine turbopumps. The approach utilizes mass and momentum source terms to model cavities and hydrodynamic blockage in inviscid, single-phase numerical calculations, reducing the computational cost of the calculations by an order of magnitude compared to traditional numerical methods. Comparison of the results from the model with experiments and high-fidelity calculations indicates agreement of the head coefficient and cavity blockage within 0.26% and 5% respectively. The computations capture rotating cavitation in a 2D inducer at the expected flow coefficient and cavitation number. The mechanism of formation and propagation of the instability is correctly reproduced.

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Experimental Investigation of the Effects of Nozzle Length on the Performance of Low Mach Number Air-Air Ejector With Entraining Diffuser

Volume 2: Aircraft Engine; Coal, Biomass and Alternative Fuels; Cycle Innovations ◽

10.1115/gt2013-94069 ◽

2013 ◽

Author(s):

A. Namet-Allah ◽

A. M. Birk

Keyword(s):

Nozzle Exit ◽

Back Pressure ◽

Flow Rates ◽

Core Flow ◽

The Core ◽

Flow Swirl ◽

Swirl Angle ◽

Mass Flow Rates ◽

Nozzle Inlet ◽

Center Body

The core flow separation in air-air ejectors is significantly affected by the length of the exhaust nozzle. This length was changed by moving the annulus’ center body end 4, 7, and 12 cm upstream and 1 cm downstream of the nozzle inlet. The velocity profiles at the nozzle exit were measured at different mass flow rates and at 10, 20 and 30 degree swirl angles. These measurements were also conducted at two annulus’ center body end positions with elliptical and square shapes, 12 and 7 cm upstream of the nozzle inlet, using two nozzle exit diameters. At 4, 7, and 12 cm upstream and 1 cm downstream of the nozzle inlet, the ejector performance was also measured at ambient temperature and at different flow swirl angles. It was found that the square shape of the annulus’ center body decreased the size of the core flow separation behind the annulus center body compared with the elliptical shape by improving the flatness of the flow velocity at the nozzle exit under different mass flow rates, swirl angles, positions of the annulus’ center body, and nozzle exit diameters. It was seen that moving the end of the annular center body upstream has considerable effects on the size and nature of the core separation behind the annulus’ center body and consequently on the ejector performance. At a zero swirl angle, the ejector pumping ratio slightly increased, decreased, and then increased again by moving the annulus’ center body from 12 cm to 7 cm upstream, from 7 cm to 4 cm upstream, and from 4 cm upstream to 1 cm downstream of the nozzle inlet respectively. These changes in the annulus’ center body position caused the back pressure coefficient to decrease, increase, and then increase again. The same trend in pumping ratio and back pressure was observed for both 10 and 20 degree flow swirl angle conditions when the annulus’ center body was moved as described.

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Computational Analysis of Cryogenic Flow Through a Control Valve

Volume 2: Symposia, Parts A, B, and C ◽

10.1115/fedsm2003-45120 ◽

2003 ◽

Cited By ~ 6

Author(s):

Russell L. Daines ◽

Jody L. Woods ◽

Peter R. Sulyma

Keyword(s):

Rocket Engine ◽

Control Valve ◽

Flow Coefficient ◽

Computational Time ◽

Working Fluid ◽

Double Precision ◽

Cryogenic Flow ◽

Flow Through ◽

Valve Modeling ◽

Verification Process

The initial efforts to develop the capability to model valves used in rocket engine component testing at Stennis Space Center are documented. An axisymmetric model of a control valve with LN2 as the working fluid was developed. The goal was to predict the effect of change in the plug/seat region of the valve prior to testing. The valve flow coefficient was predicted for a range of plug positions. Verification of the calculations was carried out to quantify the uncertainty in the numerical answer. The modeled results compared well qualitatively to experimental trends. Additionally, insights into the flow processes in the valve were obtained. Benefits from the verification process included the ability to use coarser grids and insight into ways to reduce computational time by using double precision accuracy and non-integer grid ratios. Future valve modeling activities will include shape optimization of the valve/seat region and dynamic grid modeling.

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Source Term Based Modeling of Rotating Cavitation in Turbopumps

Journal of Engineering for Gas Turbines and Power ◽

10.1115/1.4042302 ◽

2019 ◽

Vol 141 (6) ◽

Author(s):

A. G. Vermes ◽

C. Lettieri

Keyword(s):

Price Competition ◽

High Speed ◽

Large Scale ◽

Source Term ◽

Experimental Testing ◽

Rocket Engine ◽

Computational Cost ◽

Flow Coefficient ◽

Rocket Engines ◽

Launch Vehicles

The recent growth of private options in launch vehicles has substantially raised price competition in the space launch market. This has increased the need to deliver reliable launch vehicles at reduced engine development cost and has led to increased industrial interest in reduced order models. Large-scale liquid rocket engines require high-speed turbopumps to inject cryogenic propellants into the combustion chamber. These pumps can experience cavitation instabilities even when operating near design conditions. Of particular concern is rotating cavitation (RC), which is characterized by an asymmetric cavity rotating at the pump inlet, which can cause severe vibration, breaking of the pump, and loss of the mission. Despite much work in the field, there are limited guidelines to avoid RC during design and its occurrence is often assessed through costly experimental testing. This paper presents a source term based model for stability assessment of rocket engine turbopumps. The approach utilizes mass and momentum source terms to model cavities and hydrodynamic blockage in inviscid, single-phase numerical calculations, reducing the computational cost of the calculations by an order of magnitude compared to traditional numerical methods. Comparison of the results from the model with experiments and high-fidelity calculations indicates agreement of the head coefficient and cavity blockage within 0.26% and 5%, respectively. The computations capture RC in a two-dimensional (2D) inducer at the expected flow coefficient and cavitation number. The mechanism of formation and propagation of the instability is correctly reproduced.

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