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The decarbonization of production processes plays an important role on the way to environmentally friendly economy. Especially, the implementation of high temperature heat pumps (HTHP) offers a great potential to replace fossil fuel-based energy infrastructure. A major issue for the introduction of HTHP is the initial cost and regarding the payback period. However, there is still potential in increasing the coeffcient of performance (COP) of HTHP for the economic integration in existing industrial processes. One important possibility is the dedicated development and design of turbocompressors for this application and the planned heat transfer medium including the aerodynamic optimization of compressor geometry. Against this background an automated aerodynamic optimization method for radial compressor blade geometry for superheated steam is presented. The optimization refers to two different operating points of the HTHP and focuses on maximizing the isentropic efficiency of the impeller geometry as well as the pressure ratio. The algorithm is accelerated by data-driven metamodels and is implemented in a high-performance cluster environment. The boundary condition of the inherent computational fluid dynamics (CFD) calculation comes from the thermodynamic cycle calculation of the whole HTHP system. A two-stage compression with intercooling between the compressor stages are foreseen. Our approach shows an increment of both objective functions in both operating points and the satisfaction of further side conditions for the low pressure compressor (LPC). Furthermore, it results in an increment of 5 percent points of isentropic efficiency and 13 percent points of static to total pressure ratio in comparison to our initial geometry. These impeller optimizations result in a COP increment of 5 percent. The resulting geometry will be interpreted in the context of aerodynamic behavior. Based on that results additionally, a flow-cut optimization for the high pressure compressor (HPC) is given and evaluated. The results are comparable to aerodynamic optimization in different research fields like aircraft engines or stationary gas turbines and contribute to optimized multistage compressor design for HTHP.

As part of Clean Sky 2 (Engines ITD), a working group, formed by the DLR Institute of Propulsion Technology, MTU Aero Engines and GKN Aerospace, is setting up a complete compression system module (LPC/ICD/HPC). The aim of the project is to shorten the intermediate compressor duct (ICD), which connects the low pressure compressor (LPC) and high pressure compressor (HPC). This leads to a reduction of the axial length and weight of the overall engine and thus a reduction of the specific fuel consumption. In order to identify measurement-technologies and mitigate risks for this experimental set-up, a non-rotating testrig was set up in advance to investigate two different test vehicles. In a first step an ICD demonstrator was set up, which had a reduced axial length of 25%, compared to currently operating designs. This design predicted no separation or performance losses inside of the aggressive s-shaped contour. In a second step another demonstrator was installed with an even shorter axial length of 50% axial length reduction, compared to todays state of the art configurations. In this demonstrator the design was chosen to exceed the functional limit according to today’s CFD prediction and thus explore the potential for further length reduction in the future and to create a data base for numerical tool validation. The two configurations were tested at DLR Cologne. The test channel offers the possibility to use a variety of different measurement techniques to determine the flow behavior inside the ICD. This paper presents the oil streak pattern taken on the hub area between the struts and the ones taken on the pressure and suction side of the outlet guide vanes (OGV- LPC) for both demonstrators.

recognized as one of the most difficult challenges for humankind. Despite agreed climate targets, global emissions of Greenhouse gases have actually increased over the last three decades. [1] Especially in the energy-intensive industrial sector, there is still great potential for CO2-reduction, which has not yet been exploited. The Institute of Low-Carbon Industrial Processes dedicates its research efforts to these savings. The present article sheds light on the most recent and upcoming research topics for reducing or eliminating fossil energy sources as well as process-related CO2-emissions caused by industrial production.

The aerodynamic design of turbomachinery is an exciting engineering problem that involves a vast design space and requires proficiency in multiple engineering disciplines. For multi-stage compressors, the large dimension of the design space arises from the geometric freedom to shape each blade row individually. Although the basis of most designs is formed by throughflow computations, the state-of-the art for aerodynamic evaluation is steady-state RANS. The computational expense for a single evaluation of a multi-stage machine with RANS is affordable, shape optimizations with hundreds of evaluations becomes expensive. This paper presents two strategies to tackle these problems: Firstly, a novel airfoil family is employed that helps to reduce the number of design parameters. It was generated based on a parametric study of optimal airfoil shape for varying design requirements. Secondly, an optimizer is used that supports simultaneous evaluations of the objective function on multiple fidelity levels. Here, cheap throughflow calculations are conducted on the low-fidelity side and RANS is used on the high-fidelity side. Inside the optimizer, multi-fidelity support is enabled by a Co-Kriging surrogate model. This approach is expected to speed-up the optimization by guiding expensive evaluations of the 3D CFD setup into promising regions of the design space. These strategies are assessed by testing different optimization setups for a four stage compressor. Both purely low and high-fidelity optimizations are conducted as well as multi-fidelity optimizations. In comparison to the baseline design, the isentropic efficiency predicted by RANS for the working line points on the 95% and the 100% speed line is increased by above 1.2%. The driving mechanism is a load redistribution between the stages that mainly reduces the shock strength of the second rotor.

The geared turbofan aircraft engine is an efficient type of aircraft engine. By using a gearbox, the fan can run at a different rotational speed in comparison to the shaft connecting low pressure compressor and turbine. To assess the in-house tools for preliminary design, a geared turbofan, with the same requirements as the Pratt&Whitney 1000G, is designed, including thermodynamic cycle, turbine, compressor, burner and structural mechanics. The goal of this study is the design of the compression system, composed of the low-pressure compressor, the inter-compressor duct and the high-pressure compressor. The following steps have been carried out to design the compression system: first, a preliminary design of both compressors is realized, following by a detailed design, conducted with throughflow calculations and the use of a database of optimal airfoils [1]. Three stages have been chosen for the low-pressure compressor and eight stages for the high-pressure compressor. In order to reach the target conditions for all operating points from the specifications, the schedule for the variable guide vanes has also been decided on. The final design is validated with 3D CFD calculations. Additionally, the interaction of the inter-compressor duct with the low-pressure and the highpressure compressor is investigated.

The 4.5 stage DLR research compressor Rig250, which is representative of the front stages of a modern high pressure compressor, has been investigated experimentally and numerically. In order, to enable comparisons with the time-resolved experimental data both steady and unsteady numerical simulations have been performed. In these simulations, both time- and frequency-domain methods have been employed. For the time-domain calculations the model consisted of the full annulus mesh of the first two stages including the IGVs and struts located upstream in a swan neck. In total, the mesh has about 1x109nodes and is distributed over 6000 blocks. For the frequency-domain calculations the Harmonic Balance (HB) method has been used. By exploiting the spatial and temporal periodicity of the configuration the HB method only requires a single passage to be resolved per blade row, thus significantly reducing the mesh size and calculation time. In this work the results of the time- and frequency-domain simulations are compared with each other and the available steady and unsteady experimental data. In particular, the importance of selecting the appropriate Fourier modes and the challenge this presents in multistage configurations is emphasized. The draw backs and advantages of the two numerical methods are discussed., {"references":["[1] Bardina, J., Ferziger, J. H., and Rogallo, R. S., ( 1985) . Effect of rotation on isotropic turbulence: computation and modelling. Journal of Fluid Mechanics, 154, p. 321336","[2] Becker, K., Heitkamp, K., and Kügeler, E. (2010). Recent progress in a hybrid -grid CFD solver for turbomachinery flows. V European Conference on Computation Fluid Dynamics ECCOMAS CFD","[3] Franke, M., Röber, T., Kügeler, E., and Ashcroft, G. (2010). Turbulence treatment in steady and unsteady tur bomachinery flows. V European Conference on Computation Fluid Dynamics ECCOMAS CFD","[4] Kato, M., and Launder, B. E., (1993) . The modeling of turbulent flow around stationary and vibrating square cylinders. 9th Symposium on Turbulent Shear Flows, p. 10.4.110.4.6."]}

The 4.5 stage DLR research compressor Rig250, which is representative of the front stages of a modern high pressure compressor, has been investigated experimentally and numerically. In order, to enable comparisons with the timeresolved experimental data both steady and unsteady numerical simulations have been performed. In these simulations, both time- and frequency-domain methods have been employed. For the time-domain calculations the model consisted of the full annulus mesh of the first two stages including the IGVs and struts located upstream in a swan neck. In total, the mesh has about 1x109 nodes and is distributed over 6000 blocks. For the frequency-domain calculations the Harmonic Balance (HB) method has been used. By exploiting the spatial and temporal periodicity of the configuration the HB method only requires a single passage to be resolved per blade row, thus significantly reducing the mesh size and calculation time. In this work the results of the time- and frequencydomain simulations are compared with each other and the available steady and unsteady experimental data. In particular, the importance of selecting the appropriate Fourier modes and the challenge this presents in multistage configurations is emphasized. The drawbacks and advantages of the two numerical methods are discussed.

Ensuring a high degree of commonality among a range of products can dramatically decrease development costs. This paper aims to generate a highly versatile compressor airfoil family that covers most applications in the core compression system of aircraft engines and stationary gas turbines. This airfoil family is generated by filling a database with optimized airfoil shapes. The database is structured in seven dimensions, denominated as 'design requirements': blade stagger angle, pitch-chord ratio, profile area and the following design point properties: inlet Mach number, Reynolds number, streamtube contraction and aerodynamic loading. Additional constraints are imposed to ensure that feasible airfoils exist for each set of requirements. These constraints include limitations for profile area depending on inlet Mach number and limits for axial Mach number. To fill this seven-dimensional space, a large number of airfoils is generated by means of numerical optimization at discrete points in this space. The target is to find airfoil shapes that have low losses and ensure stable operation over wide incidence ranges. Design and off-design performance is evaluated with the blade-to-blade flow solver MISES. The solver is well established among industry and research and it is validated to a high degree by experiments. To verify the optimization strategy, it is tested on a set of existing compressor airfoils. The optimized geometries of four of the airfoils under investigation are found in the appendix. The database offers a wide variety of airfoils for different applications. Airfoils for sub- and supersonic inflow are covered as well as airfoils suited for placement at hub or casing. The benefit of using airfoils optimized for their specific purpose over having generic airfoil shapes is discussed as well. In future, this airfoil database will be used to study novel concepts for aircraft engines.

A multi-disciplinary numerical study has been carried out to improve the contours of a high-speed fan rotor outer-casing with respect to aerodynamic efficiency and broadband interaction noise. The assessment of the aerodynamic performance is provided by a 3D steady RANS flow solver (TRACE) and the acoustic assessment is performed by an analytical fan noise prediction model (PropNoise). Both tools have been integrated into a process chain driven by an advanced optimization algorithm (AutoOpti) which explores the design space and identifies optimal regions rapidly. The main result of the study is that a reduction of broadband noise at Approach flight conditions can be achieved with slight modifications of the casing contours (up to 2 dB), however this is counter-balanced by a decrease in rotor efficiency in Cruise flight conditions by around one point. This opposed trend is attributed to the higher incidence angle at Approach, causing large turbulence levels in the rotor wakes. Tonal interaction noise tends to increase substantially (by up to 8 dB) when broadband noise is reduced, which is related to constructive radial interferences induced by the modification of the rotor wake propagation in with casing contours.

The German Aerospace Center (DLR) has designed and tested a 1:3 scaled model S-shaped fan rotor as an example for a medium pressure ratio propulsor with potential application to future UHBR aero engines. In the present study, the attention will focus on the tip region in which the flow field is subject to complex flow phenomena and to the impact of the Sshape feature on the radial shock structure. Steady numerical simulations with DLR in-house solver TRACE as well as measurements were carried out. The casing is instrumented with ten piezoelectric static pressure transducers over the rotor pitch. Particle Image Velocimetry (PIV) is used to catch the flow velocities at three radial blade positions. All experimental data require a phase-locked ensemble averaging procedure. The results include the global performance of the compressor and detailed comparisons between simulations and measurements to validate the shock structures as well as the highly three dimensional-design S-shape fan.

In this study the end wall casing of a fan rotor is modified in an automated optimization process in order to achieve a higher efficiency and to gain surge margin. The casing end wall is allowed to take non-axisymmetric shapes by means of different parameters describing the contour. The control points defining the contour have only the possibility to move upwards or downwards in the radial direction. Given these non-axisymmetric shapes and that the simulations are conducted in the steady state, this study is a theoretical case not realizable in reality. The main aim is the understanding of the influence of end wall shaping on the flow and what the flow characteristics would be if the losses were minimized. The DLR in-house CFD code TRACE is used. The used optimization tool is AutoOpti, a DLR developed tool, which is based on a genetic algorithm speeded up by surrogate models. During the optimization two operating points are considered: One is calculated at a fixed backpressure (OP1) representing the operating point at maximum stage efficiency for the smooth casing and the other one (OP2) is an operating point near the surge limit calculated at a fixed outlet mass flow regulated by a PID controller. The optimization delivered some improvements that can be seen in the reduced losses in the tip area of the OP1 as well as in the surge margin. It resulted as well in particular surface distortion patterns on the contour responsible for a change in the static pressure distribution in the tip area. The study mainly focuses on the aerodynamic explanation of these improvements analyzed in different members.