Harald Schoenenborn

Aeroelasticity at Reversed Flow Conditions — Part II: Application to Compressor Surge

The prediction of blade loads during surge is still a challenging task. In the literature, the blade loading during surge is often referred to as “surge load,” which suggests that there is a single source of blade loading. In the second part of our paper it is shown that, in reality, the “surge load” may consist of two physically different mechanisms: the pressure shock when the pressure breaks down and aeroelastic excitation (flutter) during the blow-down phase in certain cases. This leads to a new understanding of blade loading during surge. The front block of a multistage compressor is investigated. For some points of the backflow characteristic, the quasi steady-state flow conditions are calculated using a Reynolds averaged Navier-Stokes (RANS)-solver. The flow enters at the last blade row, goes backwards through the compressor and leaves the compressor in front of the inlet guide vane. The results show a very complex flow field characterized by large recirculation regions on the suction sides of the airfoils and stagnation regions close to the trailing edges of the airfoils. Based on these steady solutions, unsteady calculations are performed with a linearized aeroelasticity code. It can be shown that some of the rotor stages are aerodynamically unstable in the first torsional mode. Thus, in addition to the pressure shock, the blades may be excited by flutter during the surge blow-down phase. In spite of the short blow-down phase typical for aero-engine high pressure compressors, this may lead to very high blade stresses due to high aeroelastic excitation at these special flow conditions. The analytical results compare very well with the observations during rig testing. The correct nodal diameter of the blade vibration is reproduced and the growth rate of the blade vibration is predicted quite well, as a comparison with tip-timing measurements shows. A new flutter region in the compressor map was experimentally and analytically detected.

Determination of Blade-Alone Frequencies of a Blisk for Mistuning Analysis Based on Optical Measurements

Meanwhile the importance of mistuning especially for blisks is well known. Most of the mistuning studies so far done are based on assumed statistical distributions of the eigenfrequencies. But it is important to know the real eigenfrequency distribution of the blades of blisks as they come out of the manufacturing process or after they are in service for some time and have experienced erosion, wear and FOD. Some of the current analytical procedures for mistuning calculations need the blade-alone frequencies, such as the reduced-order code Turbo-Reduce which is used in the subsequent analysis. One way to determine the eigenfrequency distribution is to ping-test every blade while damping all other blades. This procedure is very tedious and may take several days, depending on the number of blades. Further, it is very difficult to eliminate the influence of the disk and to obtain the pure blade-alone eigenfrequencies. In this paper a method for the analytical determination of the eigenfrequency distribution due to geometric imperfections based on optical measurements is described. Starting from the point-cloud from the optical measurements a procedure to obtain a FEM-Model of all blades of a blisk is presented. As the procedure can partly be run as a batch-job, the time to determine the eigenfrequency distribution is drastically reduced compared to the experimental way. The procedure is applied to a real blisk and the calculated eigenfrequencies are compared with the measured frequencies. Then mistuning calculations are performed based on this eigenfrequency distribution.Copyright

Comparison of Non-Linear and Linearized CFD Analysis of the Stator-Rotor Interaction of a Compressor Stage

The prediction of resonance amplitudes due to stator-rotor interactions is still an important task within the design process of turbomachinery bladings.In this paper the stator-rotor interaction of a compressor stage which consists of an inlet guide vane and a rotor blade is studied with a non-linear and a linearized CFD code. First, a quasi-3D-study of a section close to the tip region is considered. The passing of the wake of the inlet guide vane over the rotor is studied for six different vibration mode shapes of increasing complexity (first bending mode up to 4th chordwise bending mode).Whereas for low rotor speeds the comparison between linearized and non-linear calculations is quite good, large differences are found for high rotor speeds. It is shown that an acoustic interaction between the two stages with a cut-on mode is the cause for the large differences, leading to much higher unsteady pressure amplitudes on the rotor blade. This in turn leads to different aerodynamic work on the rotor blade for the different mode shapes.The extension of the investigations to 3D shows essentially the same effects.Copyright

Aeroelasticity at Reversed Flow Conditions—Part III: Reduction of Surge Loads by Means of Intentional Mistuning

Compressor surge consists of four phases: (i) pressure rise, (ii) flow breakdown, (iii) blow-down, and (iv) flow recovery. During the blow-down phase reversed flow conditions exist, where a blade may accumulate hundreds of vibration cycles, depending on the surge volume and the vibration frequency. High vibration amplitudes and blade damages were observed in the past. In Part I (GT2011-45034) a compressor cascade was analyzed experimentally and analytically at steady reversed flow conditions. It has been shown that (i) the steady flow field can be predicted well by CFD analysis, (ii) the overall damping coefficient calculated by unsteady CFD compares reasonably well with measurements, and (iii) a blade may become unstable at certain reversed flow conditions. In Part II (GT2011-45035) the analytical procedures used in Part I were applied to the front part of a multistage HPC for reversed flow conditions. It was found that surge loads consist in reality of two physically different phenomena (i) the pressure wave during the flow breakdown leading to rather low blade stresses and (ii) flutter during the blow-down phase which may lead to very high blade stresses and damages during surge for some stages. As it is well known that intentional mistuning is a way to mitigate flutter, intentional mistuning is investigated in Part III of the paper at reversed flow conditions. At first, a CFD study of a single airfoil is presented showing the dependency of aerodynamic damping upon flow angle and pressure ratio over the airfoil at reversed flow conditions, including intentional mistuning studies. Secondly, an investigation is presented which shows experimentally and analytically that surge stresses can be reduced significantly by the use of intentional mistuning. In a multistage compressor test rig, one rotor stage, which experienced very high stresses during surge, was subjected to a cutback on every second blade, leading to significantly reduced surge stresses. Analytically, an aeroelastic eigenvalue analysis showed the same behavior.

Contribution to Free and Forced Vibration Analysis of an Intentionally Mistuned Blisk

Most of the experimental mistuning studies are performed using a blisk with random mistuning only. Intentional mistuning is often investigated analytically with respect to aeroelasticity, as it is well known that intentional mistuning reduces the flutter risk due to less interaction between the blades.In this paper, an intentionally mistuned test blisk is investigated both analytically and experimentally with respect to free and forced vibrations. First, free vibrations are studied and aliasing effects for the intentionally mistuned blisk are analyzed in comparison with a tuned blisk. A comparison between the experimentally determined dominant nodal diameters and the computed ones shows good agreement.Then, the blisk is experimentally excited by a travelling wave for various engine orders. Similar investigations are performed with a FEM model of the blisk and a reduced-order code. The amplification factor for some modes and several blisks is compared. The influence of the disc onto the blade mode shapes is studied for the tuned and mistuned case without and with aerodynamic coupling effects.Cyclic spacing of vanes is a concept to reduce the vibration level of downstream rotor blades by distributing the excitation onto more engine orders while reducing the overall excitation level. In this paper it is shown for blisks with and without intentional mistuning that care should be taken in applying this concept in the vicinity of veering regions, because the amplification factor in a veering region may become much higher than compared to other nodal diameters.Copyright

Aeroelasticity at Reversed Flow Conditions – Part 1: Numerical and Experimental Investigations of a Compressor Cascade with Controlled Vibration

The prediction of flutter and forced response at normal flow conditions has become a standard procedure during the design of compressor airfoils. But at severe off-design conditions, the flow field becomes very complex, especially during the surge blow-down phase where reversed flow conditions occur. The correct prediction of the unsteady pressures and the resulting aerodynamic excitation or damping at these conditions remains an extremely challenging task. In the first part of the paper, basic investigations for these flow conditions are presented. Aeroelastic calculations during compressor surge are shown in the second part. Experimental investigations were performed in the Annular Test Facility for non-rotating cascades at EPF Lausanne. The test cascade was exposed to flow conditions as expected during the surge blow-down phase which is characterized by large separation regions. Measurements of the steady-state flow conditions on the blade surface, at the outer wall, upstream and downstream of the cascade provided detailed information about the steady flow conditions. The cascade was then subjected to controlled vibration of the blades with constant amplitudes and inter-blade phase angles. Unsteady pressure measurements on the blade surface and at the casing wall provided information about the resulting unsteady flow conditions. Analytical CFD calculations were performed. The steady flow field was calculated using a RANS code. Based on the steady-state flow field, unsteady calculations applying a linearized code were carried out. The agreement between measurements and calculations shows that the steady flow as well as the unsteady flow phenomena can be predicted quantitatively. In addition, knowing the blade vibration mode shape, which in this case is a torsion mode, the aerodynamic damping can be determined for the corresponding flow conditions.

Experimental and Analytical Mistuning Analysis of a Blisk at Lab Conditions and Under Rig Conditions Using Tip Timing

Determination of the amplification factor due to mistuning is an important task for the safe design of turbomachinery, specially for blisk-design with low mechanical damping. The complexity of the environment effects increases from measurements in the laboratory to measurements in test rigs and engines. Also the uncertainties of the boundary conditions increase. In this paper measurements of the amplification factors due to mistuning at various conditions are presented, starting from simple lab measurements up to rig measurements. Calculations with a Reduced-Order Code based on measured frequency distributions and FEM are performed and compared with the measurements. For the determination of the amplification factor a mistuning rig with travelling wave excitation was built. For a small demo-blisk the amplification factor was determined. Then a real blisk was tested. Afterward, the same blisk was built into a two-stage axial compressor rig. Here, tip-timing measurements were performed under rig conditions (centrifugal force, flow conditions). As tip-timing measures the vibration amplitude of each single airfoil, an amplification factor can be determined for each resonance.Copyright

Aeroelasticity at Reversed Flow Conditions: Part 2—Application to Compressor Surge

The prediction of blade loads during surge is still a challenging task. In literature the blade loading during surge is often referred to as “surge load”, which suggests that there is a single source of blade loading. In the second part of the paper it is shown that the “surge load” in reality may consist of two physically different mechanisms: the pressure shock when the pressure breaks down and aeroelastic excitation (flutter) during the blow-down phase in certain cases. This leads to a new understanding of blade loading during surge. The front block of a multistage compressor is investigated. For some points of the backflow characteristic the quasi steady-state flow conditions are calculated using a RANSsolver. The flow enters at the last blade row, goes backwards through the compressor and leaves the compressor in front of the inlet guide vane. The results show a very complex flow field characterized by large recirculation regions on the suction sides of the airfoils and stagnation regions close to the trailing edges of the airfoils. Based on these steady solutions unsteady calculations are performed with a linearized aeroelasticity code. It can be shown that some of the rotor stages are aerodynamically unstable in the first torsional mode. Thus, in addition to the pressure shock the blades may be excited by flutter during the surge blow-down phase. In spite of the short blow-down phase typical for aero-engine high pressure compressors, this may lead to very high blade stresses due to high aeroelastic excitation at these special flow conditions. The analytical results compare very well with the observations during rig testing. The correct nodal diameter of the blade vibration is reproduced and the growth rate of the blade vibration is predicted quite well, as a comparison with tip-timing measurements shows. A new flutter region in the compressor map was detected experimentally and analytically. NOMENCLATURE A Amplitude f Blade eigenfrequency IBPA Inter-blade phase angle LE Leading edge ND Nodal diameter p Pressure Ref Reference S/G Strain gauge stat Static T Vibration period TE Trailing edge t Time tot total δ Aerodynamic damping (log. decrement) 1F First bending vibration mode 1T First torsional vibration mode

Analysis of the Effect of Multi-Row and Multi-Passage Aerodynamic Interaction on the Forced Response Variation in a Compressor Configuration: Part 1 — Aerodynamic Excitation

The aeroelastic prediction of blade forcing is still a very important topic in turbomachinery design. Usually, the wake from an upstream airfoil and the potential field from a downstream airfoil are considered as the main disturbances. In recent years, it became evident that in addition to those two mechanisms Tyler-Sofrin modes, also called scattered or spinning modes, may have a significant impact on blade forcing.In Schrape et al. [9] it was found that in multi-row configurations not only the next, but also the next but one blade row is very important as it may create a large circumferential forcing variation which is fixed in the rotating frame of reference.In the present paper a study of these effects is performed on the basis of a quasi 3D multi-row and multi-passage compressor configuration. For the analysis a harmonic balancing code, which was developed by DLR Cologne, is used for various setups and the results are compared to full-annulus unsteady calculations. It is shown that the effect of the circumferentially different blade excitation is mainly contributed by the Tyler-Sofrin modes and not to blade-to-blade variation in the steady flow field.The influence of various clocking positions, coupling schemes and number of harmonics onto the forcing is investigated. It is also shown that along a speed-line in the compressor map the blade-to-blade forcing variation may change significantly.In addition, multi-row flutter calculations are performed, showing the influence of the upstream and downstream blade row onto aerodynamic damping.The effect of these forcing variations onto random mistuning effects is investigated in the second part of the paper.Copyright

Consideration of Interface Damping in Shrouded Mistuned Turbine Blades

Running turbines are exposed to high mechanical load. Due to gas excitations the structure can vibrate with high oscillation amplitudes which can damage the turbine blades. Mistuning can additionally lead to high local stresses which must be taken into account in the turbine design process. Introducing damping due to friction in the interface of shrouded turbines can be used to decrease this oscillation amplitudes. The computation of full turbine finite-element models with nonlinear coupling forces causes high computational costs. As a consequence, Component Mode Synthesis methods are used to reduce the number of DOFs of each blade substructure. Mistuning of the blades can now be applied in modal space. Coupling of the mistuned substructures is done by nonlinear interface forces which have to be included in the substructuring formulation. The resulting reduced and mistuned system with nonlinear coupling forces is solved with a Harmonic Balance Method such that the effect of mistuning and interface damping can be studied very efficiently.