Sensitivity of Supersonic ORC Turbine Injector Designs to Fluctuating Operating Conditions

The design of an efficient organic rankine cycle (ORC) expander needs to take properly into account strong real gas effects that may occur in given ranges of operating conditions, which can also be highly variable. In this work, we first design ORC turbine geometries by means of a fast 2-D design procedure based on the method of characteristics (MOC) for supersonic nozzles characterized by strong real gas effects. Thanks to a geometric post-processing procedure, the resulting nozzle shape is then adapted to generate an axial ORC blade vane geometry. Subsequently, the impact of uncertain operating conditions on turbine design is investigated by coupling the MOC algorithm with a Probabilistic Collocation Method (PCM) algorithm. Besides, the injector geometry generated at nominal operating conditions is simulated by means of an in-house CFD solver. The code is coupled to the PCM algorithm and a performance sensitivity analysis, in terms of adiabatic efficiency and power output, to variations of the operating conditions is carried out.

On the Application of Frequency-Domain Methods to Multistage Turbomachinery

Due to the relative motion between adjacent blade rows the aerodynamic flow fields within turbomachinery are normally dominated by deterministic, periodic phenomena. In the numerical simulation of such unsteady flows (nonlinear) frequency-domain methods are therefore attractive as they are capable of fully exploiting the given spatial and temporal periodicity, as well as capturing or modelling flow nonlinearity. Central to the efficiency and accuracy of such frequency-domain methods is the selection of the frequencies and the circumferential modes to be resolved in simulations. Whilst trivial in the context of the simulation of a single compressor- or turbine-stage, the choice of solution modes becomes substantially more involved in multi-stage configurations. In this work the importance of mode scattering, in the context of the unsteady aerodynamic field, is investigated and quantified. It is shown that scattered modes can substantially impact the unsteady flow field and are essential for the accurate modelling of wake propagation within multistage configurations. Furthermore, an iterative approach is outlined, based on the spectral analysis of the circumferential modes at the interfaces between blade rows, to identify the dominant solution modes that should be resolved in the adjacent blade row. To demonstrate the importance of mode scattering and validate the approach for their identification the unsteady blade row interaction within a 4.5 stage axial compressor is computed using both the harmonic balance method and, based on a full annulus midspan simulation, a time-domain method. Through the inclusion of scattered modes it is shown that the solution quality of the harmonic balance results is comparable to that of the nonlinear time-domain simulation.

Study of Semi-Inverse Problem for the Design of Annular S-Shaped Compressor Transition Duct

A new aerodynamic design strategy of the S-shaped transition duct between two compressor components was studied. Based on the controlled wall pressure gradient distribution and the wall velocity distribution, a semi-inverse problem of the transition duct was proposed, the corresponding inverse and direct approach codes were developed. To verify the feasibility of this method, two axial-centrifugal compressor transition ducts were designed. The results show that the static pressure distribution on the inner wall and the duct geometry both can be controlled freely by adjusting the inverse design parameters. The designed inner wall pressure distribution can be realized through a numerical matching procedure of the outer wall geometry based on the direct problem. The new design method is practicable that, without searching the optimal solution of the static pressure distribution of the inner wall, the total pressure coefficient can be at least 0.92.

Design and Test of a Transonic Axial Splittered Rotor

A new design procedure was developed that uses commercial-off-the-shelf software (MATLAB, SolidWorks, and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial compressor rotor with splitter blades. Predictive numerical simulations were conducted and experimental data were collected in a Transonic Compressor Rig. This study advanced the understanding of splitter blade geometry, placement, and performance benefits. In particular, it was determined that moving the splitter blade forward in the passage between the main blades, which was a departure from the trends demonstrated in the few available previous transonic axial compressor splitter blade studies, increased the mass flow range with no loss in overall performance. With a large 0.91 mm (0.036 in) tip clearance, to preserve the integrity of the rotor, the experimentally measured peak total-to-total pressure ratio was 1.69 and the peak total-to-total isentropic efficiency was 72 percent at 100 percent design speed. Additionally, a higher than predicted 7.5 percent mass flow rate range was experimentally measured, which would make for easier engine control if this concept were to be included in an actual gas turbine engine.

Design Method and Performance Effects of Curvature-Smooth Centrifugal Compressor Blades

Blade curvature has an important effect on the pressure distribution over the blade surface as well as the quality of airfoil machining, but curvature smoothing method in design environments is little studied. A new geometry generator is developed for the design of curvature-smooth airfoils. The sections of an airfoil can be initially designed by a conventional method, and each of the sections is then regenerated by fitting a fifth order B-spline curve. The first order differential of the curvature with respect to arc length of the blade surface is used as a curvature smoothness measure. The distance between discrete points and the curve, and the integral of the square of the first derivative along the streamwise section curve are simultaneously minimized in the fitting process. The design of a transonic centrifugal compressor for turbocharger application is presented to study the effects of the curvature-smooth blade on centrifugal compressor performance. A numerical study was carried out and the results are discussed.

Assessment of the Indirect Combustion Noise Generated in a Transonic High-Pressure Turbine Stage

Indirect combustion noise, generated by the acceleration and distortion of entropy waves through the turbine stages, has been shown to be the dominant noise source of gas turbines at low-frequencies and to impact the thermoacoustic behavior of the combustor. In the present work, indirect combustion noise generation is evaluated in the realistic, fully 3D transonic high-pressure turbine stage MT1 using Large-Eddy Simulations (LES). An analysis of the basic flow and the different turbine noise generation mechanisms is performed for two configurations: one with a steady inflow and a second with a pulsed inlet, where a plane entropy wave train at a given frequency is injected before propagating across the stage generating indirect noise. The noise is evaluated through the Dynamic Mode Decomposition of the flow field. It is compared with previous 2D simulations of a similar stator/rotor configuration, as well as with the compact theory of Cumpsty and Marble. Results show that the upstream propagating entropy noise is reduced due to the choked turbine nozzle guide vane. Downstream acoustic waves are found to be of similar strength to the 2D case, highlighting the potential impact of indirect combustion noise on the overall noise signature of the engine.

Effects of Flow Distortions as They Occur in S-Duct Inlets on the Performance and Stability of a Jet Engine

One of the research areas at the Institute of Jet Propulsion focuses on the design and optimization of s-shaped engine inlet configurations. The distortion being evoked within such inlet ducts should be limited to ensure an optimal performance, stability, and durability of the engine’s compression system. Computational Fluid Dynamics (CFD) play a major role in the design process of bent engine inlet ducts. The flow within such ducts can be computed, distortion patterns can be visualized, and related distortion coefficients are easily calculated. The impact of a distortion on flow phenomena within the compressor system can, however, only be computed with major computational efforts and thus the quality of an s-duct design in development is usually assessed by analyzing the evoked distortion with suitable distortion coefficients without a true knowledge of the duct’s influence on the downstream propulsion system. The influence of inlet distortion on both the performance and stability of the Larzac 04 jet engine was parameterized during experimental investigations at the engine test bed of the Institute of Jet Propulsion. Both pressure and swirl distortion patterns as they typically occur in s-duct inlet configurations were reproduced with distortion generators. Pressure distortion patterns were generated using seven types of distortion screens. The intensity of the distortion varies with the mesh size of the screen whereas the extension of the distortion is defined by the dimensions of the screen in radial and circumferential direction. A typical counter rotating twin-swirl was generated with a deltawing installed upstream of the compressor system. First, the development of flow distortion was analyzed for several engine operating points (EOPs). A linear relation between the total pressure loss in the engine inlet and the EOPs was found. Secondly, the flow within the compressor system with an inlet distortion was analyzed and unsteady flow phenomena were detected for severe inlet distortions. Finally, the effect of both pressure and swirl distortion on the performance and stability of the test vehicle was parameterized. A loss in engine performance with increasing inlet distortion is observable. The limiting inlet distortion with respect to engine stability was found and moreover it was shown that pressure distortion has a stronger influence on the stability of the compressor system compared to a counter rotating twin-swirl distortion. The presented parameterization is essential for the s-duct design, which is under development for an experimental set-up with the Larzac 04 jet engine.

The Quantification of Impact of Geometrical and Operational Parameters Uncertainty on the Numerical Simulation Results of Axial Turbine Working Process

The quantification of geometric and physical variables uncertainty impact on turbomachinery row workflow was conducted using several untwisted airfoil cascades of axial turbine nozzle blades with uniform section throughout the channel height. Profile loss coefficient, mass flow parameter, outlet flow angle were accepted as controlled performance criteria. The series of computational calculations were carried out for these cascades. The first group of calculations was aimed at the identification of the impact of geometric parameters uncertainty on nozzle blades parameters. The second group of calculations was conducted for identifying the studied parameters depending on the flow parameters changes that are used as boundary conditions in the simulation. The obtained results showed that initial data uncertainty in CFD calculations has a significant impact on the obtained quantitative estimates. The difference between calculated data modified in accordance with the geometry technological tolerances and workflow parameters measurement error may exceed 5% by value of the considered criteria.

Simulations of Unsteady Blade Row Interactions Using Linear and Non-Linear Frequency Domain Methods

This paper compares various approaches to simulate unsteady blade row interactions in turbomachinery. Unsteady simulations of turbomachinery flows have gained importance over the last years since increasing computing power allows the user to consider 3D unsteady flows for industrially relevant configurations. Furthermore, for turbomachinery flows, the last two decades have seen considerable efforts in developing adequate CFD methods which exploit the rotational symmetries of blade rows and are therefore up to several orders of magnitude more efficient than the standard unsteady approach for full wheel configurations. This paper focusses on the harmonic balance method which has been developed recently by the authors. The system of equations as well as the iterative solver are formulated in the frequency domain. The aim of this paper is to compare the harmonic balance method with the time-linearized as well as the non-linear unsteady approach. For the latter the unsteady flow fields in a fan stage are compared to reference results obtained with a highly resolved unsteady simulation. Moreover the amplitudes of the acoustic modes which are due to the rotor stator interaction are compared to measurement data available for this fan stage. The harmonic balance results for different sets of harmonics in the blade rows are used to explain the minor discrepancies between the time-linearized and unsteady results published by the authors in previous publications. The results show that the differences are primarily due to the neglection of the two-way coupling in the time-linearized simulations.