Research Papers: Gas Turbines: Combustion, Fuels, and Emissions

Characterization of a Supersonic Turbine Downstream of a Rotating Detonation Combustor

[+] Author and Article Information
Z. Liu

School of Mechanical Engineering,
Purdue University,
West Lafayette, IN 47906
e-mail: liu1752@purdue.edu

J. Braun

School of Mechanical Engineering,
Purdue University,
West Lafayette, IN 47906
e-mail: jamesbraun91@gmail.com

G. Paniagua

School of Mechanical Engineering,
Purdue University,
West Lafayette, IN 47906
e-mail: gpaniagua@me.com

1Corresponding author.

Contributed by the Aircraft Engine Committee of ASME for publication in the JOURNAL OF ENGINEERING FOR GAS TURBINES AND POWER. Manuscript received June 23, 2018; final manuscript received July 1, 2018; published online October 4, 2018. Editor: Jerzy T. Sawicki.

J. Eng. Gas Turbines Power 141(3), 031501 (Oct 04, 2018) (13 pages) Paper No: GTP-18-1324; doi: 10.1115/1.4040815 History: Received June 23, 2018; Revised July 01, 2018

Rotating detonation combustors (RDCs) offer theoretically a significant total pressure increase, which may result in enhanced cycle efficiency. The fluctuating exhaust of RDC, however, induces low supersonic flow and large flow angle fluctuations at several kHz, which affects the performance of the downstream turbine. In this paper, a numerical methodology is proposed to characterize a supersonic turbine exposed to fluctuations from RDC without any dilution. The inlet conditions of the turbine were extracted from a three-dimensional (3D) unsteady Reynolds-averaged Navier–Stokes simulation of a nozzle attached to a rotating detonation combustor, optimized for minimum flow fluctuations and a mass-flow averaged Mach number of 2 at the nozzle outlet. In a first step, a supersonic turbine able to handle steady Mach 2 inflow was designed based on a method of characteristics solver and total pressure loss was assessed. Afterward, unsteady simulations of eight stator passages exposed to periodic oblique shocks were performed. Total pressure loss was evaluated for several oblique shock frequencies and amplitudes. The unsteady stator outlet profile was extracted and used as inlet condition for the unsteady rotor simulations. Finally, a full stage unsteady simulation was performed to characterize the flow field across the entire turbine stage. Power extraction, airfoil base pressure, and total pressure losses were assessed, which enabled the estimation of the loss mechanisms in supersonic turbine exposed to large unsteady inlet conditions.

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Fig. 1

(a) Geometry of the RDC and nozzle and (b) 3D optimized nozzle geometry

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Fig. 2

(a) Instantaneous static temperature contour of the 2D RDC. Nozzle outlet profile: (b) static pressure, (c) static temperature, (d) Mach number, and (e) flow angle.

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Fig. 3

Blade to blade view of the (a) supersonic stator and (b) supersonic rotor

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Fig. 4

(a) Mach contour of the Mach 5 validation case, (b) comparison of skin friction, and (c) wall static pressure

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Fig. 5

Sod shock tube problem: (a) density field after the release of the diaphragm [23] and (b) density variation along the axial position [23]

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Fig. 6

(a) Meridional view of the stage channel, (b) numerical grid with a close up of the stator trailing edge

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Fig. 7

(a) Rotor relative total pressure versus grid size. (b) Torque in function of the convergence. Contour of relative total pressure at rotor outlet for the (c) medium mesh and (d) fine mesh. (e) Local discretization error of the medium mesh.

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Fig. 8

Steady flow field of the supersonic turbine at midspan for an inlet Mach number of 2

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Fig. 10

Instantaneous contour at stator outlet: (a) Mach number, (b) total pressure, and (c) total temperature (f¯ = 0.24, A¯ = 1)

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Fig. 9

Unsteady Reynolds-averaged Navier–Stokes stator simulation at midspan exposed to oblique shocks for three consecutive phases (f¯ = 0.24, A¯ = 1): (a) Mach contour and (b) total pressure contour. (c) Close up of the instantaneous shock patterns in (a) during phase 3.

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Fig. 12

(a) Instantaneous flow field of the stator base region (f¯ = 0.24, A¯ = 1) and (b) stator outlet and base static pressure as a function of time

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Fig. 11

Instantaneous stator fluctuation as a function of the azimuthal angle: (a) flow angle, (b) static pressure, (c) Mach number, and (d) static temperature (f¯ = 0.24, A¯ = 1)

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Fig. 17

Instantaneous rotor fluctuations as a function of the azimuthal angle: (a) relative flow angle, (b) static pressure, (c) relative Mach number, and (d) static temperature (f¯= 0.24, A¯ = 1)

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Fig. 16

Instantaneous contour of rotor outlet: (a) relative Mach number, (b) relative total pressure, and (c) relative total temperature (f¯ = 0.24, A¯ = 1)

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Fig. 15

(a) Flow field of a URANS rotor simulation at midspan for three phases (f¯ = 0.24, A¯ = 1). (b) Close up of the instantaneous shock patterns at the close-up position in (a) during phase3.

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Fig. 14

Stator total pressure loss as function of reduced frequency and nondimensional amplitude

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Fig. 13

Instantaneous Mach contour and instantaneous Mach number fluctuations at stator outlet as a function of the azimuthal angle at: (a) f¯ = 0.24, A¯ = 0.5, (b) f¯ = 0.24, A¯ = 1, and (c) f¯= 0.23, A¯ = 2

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Fig. 18

(a) Flow field of a URANS stage simulation at midspan exposed to oblique shocks for three phases. (b) Close up on the instantaneous shock patterns at the location in (a) during phase 2. (c) 3D flow structures within the stage during phase 2. (d) Static temperature of stator and rotor blades for three phases (f¯ = 0.24, A¯ = 1).

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Fig. 20

(a) Instantaneous flow field of the URANS stage simulation at midspan (f¯ = 0.24, A¯ = 1). (b) Frequency spectrum of static pressure at the stator inlet and rotor inlet.

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Fig. 19

Instantaneous stage fluctuations as a function of the azimuthal angle: (a) flow angle, (b) static pressure, (c) Mach number, and (d) static temperature (f¯ = 0.24, A¯ = 1)



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