The objective of this study is to quantify the sensitivity of blade roughness on the overall performance of a 10-stage high-pressure compressor of the jet engine type V2500-A1. The Reynolds-Aver-aged Navier-Stokes flow solver TRACE is used to study the multi-stage compressor. The three-dimensional numerical setup contains all geometric and aerodynamic features such as bleed ports and the variable stator vanes system. In order to estimate the effect of stage roughness on overall compressor performance, compressor maps of the CFD-model are created by modeling the surface rough- ness separately for a single stage and combinations of stages. The surface roughness values are applied to the blade's suction side of the first, center and last stage in the CFD-model by setting an equivalent sand-grain value. This equivalent sand-grain roughness is determined from non-intrusive measurements of blade surfaces from an equivalent real aircraft engine for the first, center and last stage. In addition, further simulations are conducted to analyze the performance drop of a fully rough HPC due to surface roughness. The studies are performed at the operating conditions 'cruise' and 'take-off' to cover two different Reynolds number regimes. The results show that the models with roughness in a single stage already lead to significantly lower mass flow rates because of higher block-age compared to the smooth compressor. In fact, roughness at the first stage has the biggest effect on the overall performance with a drop in performance of about 0.1% while the effect of the last stage is the smallest. This behavior is mainly caused by enhanced instabilities through the compressor changing the stage-by-stage matching of the stages downstream. In addition to the displacement of the compressor maps to a lower mass flow, a reduction of stall and choke margins is noticeable.

Deterioration of axial compressors is in general a major concern in aircraft engine maintenance. Among other effects, roughness in high-pressure compressor reduces the pressure rise and thus efficiency, thereby increasing the specific fuel consumption of an engine. Therefore, it is important to improve the understanding of roughness on compressor blading and their impact on compressor performance. To investigate the surface roughness of rotor blades of a compressors, different stages of an axial high-pressure compressor and a first-stage blisk (BLade–Integrated–dISK) of a regional aircraft engine is measured by a three-dimensional laser scanning microscope. Fundamental types of roughness structures can be identified: impacts in different sizes, depositions as isotropically distributed single elements with steep flanks and anisotropic roughness structures direct approximately normal to the flow direction. To characterise and quantify the roughness structures in more detail, roughness parameters were determined from the measured surfaces. The quantification showed that the roughness height varies through the compressor depending on the stage, position and the blade side. Overall complex roughness structures of different shape, height and size are detected regardless of the type of the blades.

The multidisciplinary performance of a bypass jet propulsion engine is investigated and simulated in this paper by using the pseudo bond graph approach. This theory is applied in the in-house software tool ASTOR (AircraftEngine Simulation of Transient Operation Research) to simulate the different operating points and the system dynamics in transient load cases with high accuracy. In this paper, the modulation of the V2500-Al turbofan jet engine in the pseudo bond graph notation is described in detail. The thermodynamic cycle of a steady-state operating point is calculated and a transient load case is investigated to simulate the system performance. The results correspond with the data from a commercial software and with literature. However, some significant discrepancies are detected, which can be explained by the current model simplification. In a nutshell, ASTOR is able to simulate the dynamic performance of the overall system and can be used to improve the process of the construction or regeneration of a complex capital good such as a jet engine.

In this paper a comparison of the performance parameters and flow characteristics of a mature commercial high bypass engine's 9-stages high pressure compressor (HPC) with steadystate mixing-plane (RANS) and unsteady RANS (URANS) CFD is carried out. The investigation is based on a numerical model of the CFM56-5C engine's HPC which is installed on the Airbus A340-300 aircraft. The compressor under investigation features the so called 3D blading which is the first of two performance improvement packages available. An experimental engine of this type equipped with extensive additional instrumentation is in use by Lufthansa Technik to deliver detailed experimental data of the gas path. Experimental results have been discussed in previous works of the authors. In order to provide long-term forecasts of an engine's state the aforementioned CFD model has been included into a multilevel engine model. To be able to evaluate the significance and applicability of the CFD results into such model, it is of great interest to which extent and level of detail it can deliver accurate performance predictions. From the comparison of both steady-state and unsteady simulation results it is found that overall compressor performance only differs negligibly while stage performance can differ significantly. It is depicted that among the stator vanes of the front stages local supersonic flow and flow separation can occur. These are not captured by the steady-state simulation to the same degree a time resolved simulation does. In rear stages differences fade and unsteady methods tend to predict better stage performance which may be due to favourable effects of rotor-stator interaction.

The system performance of a single-shaft turbojet engine is modelled with the pseudo bond graph approach in this paper. This theory is implemented in the in-house software tool ASTOR (AircraftEngine Simulation of Transient Operation Research) to simulate the overall dynamic of the turbojet model engine P200SX. In ASTOR, the transient performance is calculated with dynamic and individual control volumes to determine the three conservation equations.

The decreasing performance of jet engines during operation is a major concern for airlines and maintenance companies. Among other effects, the erosion of high-pressure compressor (HPC) blades is a critical one and leads to a changed aerodynamic behavior, and therefore to a change in performance. The maintenance of BLISKs (blade-integrated-disks) is especially challenging because the blade arrangement cannot be changed and individual blades cannot be replaced. Thus, coupled deteriorated blades have a complex aerodynamic behavior which can have a stronger influence on compressor performance than a conventional HPC. To ensure effective maintenance for BLISKs, the impact of coupled misshaped blades are the key factor. The present study addresses these effects on the aerodynamic performance of a first-stage BLISK of a high-pressure compressor. Therefore, a design of experiments (DoE) is done to identify the geometric properties which lead to a reduction in performance. It is shown that the effect of coupled variances is dependent on the operating point. Based on the DoE analysis, the thickness-related parameters, the stagger angle, and the max. profile camber as coupled parameters are identified as the most important parameters for all operating points.

The overall efficiency and operational behavior of aircraft engines are influenced by the surface finish of the airfoils. During operation, the surface roughness significantly increases due to erosion and deposition processes. The aim of this study is to analyze the influence of roughness on the aerodynamics of the low-pressure turbine (LPT) of a midsized high bypass turbofan. In order to gain a better insight into the operational roughness structures, a sample of new, used, cleaned, and reworked turbine blades and vanes are measured using the confocal laser scanning microscopy technique. The measurement results show local inhomogeneities. The roughness distributions measured are then converted into their equivalent sand grain roughness ks,eq to permit an evaluation of the impact on aerodynamic losses. The numerical study is performed using the computational fluid dynamics (CFD)-solver turbomachinery research aerodynamics computational environment (TRACE) which was validated before with the existing data from rig experiments. It is observed that the influence of the surface roughness on the turbine efficiency is significant at take-off but negligible at cruise. A detailed analysis on the aerodynamics at take-off shows that very rough airfoils lead to higher profile and secondary loss. Due to the higher disturbances present in flows circulating over rough walls, the transition occurs earlier, and the momentum thickness increases in the turbulent boundary layer. The service-induced roughness structures cause an efficiency drop in the LPT of ηT=−0.16% compared to new parts. A gas path analysis showed that this results in an increased fuel flow of Δm˙f=+0.06% and an exhaust gas temperature (EGT) rise of ΔEGT=+1.2K for fixed engine pressure ratio which is equivalent to roughly 4% of the typical EGT margin of a fully refurbished engine. This result stresses the importance of roughness-induced loss in LPTs.

The Institute of Jet Propulsion and Turbomachinery of the TU Braunschweig owns a jet engine of the type V2500-A1 from the International Aero Engines AG. To conduct research on the jet engine and its components, computer models are necessary. In this paper, the reverse engineering process of the high pressure compressor (HPC) regarding its aerodynamics is presented. Thereby, the reverse engineering process starts from digitizing newly manufactured airfoils, followed by FEM-calculations to enforce the operating forces on the geometries. A computational fluid dynamics (CFD) model using these geometries is set up, considering all relevant geometric and aerodynamic features such as bleed ports and the variable stator vane (VSV) system. Using this CFD-model, the compressor map is calculated and afterwards validated by available manufacturing data [18] and by the institute's jet engine's test cell data. Because this jet engine is a highly operated and deteriorated one, a map scaling is necessary before comparing the CFD-model with the test cell data. Nevertheless, an adequate agreement of the operating behavior between scaled compressor map and test cell data is shown. To estimate the deterioration level of the jet engine's compressor and to evaluate the used scaling factors, the tip gaps inside the CFD-model were doubled and the compressor behavior was simulated. The observed effect of reduced compressor capacity and efficiency is in accordance with literature but is not able to explain the amount of the scaling factors completely.

Performance degradation due to wear of high pressure compressors (HPC) is a major concern in aero-engine operation and maintenance. Among other effects especially erosion of airfoils leads to changed aerodynamic behavior and therefore to deterioration. These affects engine performance parameter like thrust specific fuel consumption (TSFC) and exhaust gas temperature (EGT). Reaching EGT-limit, the engine typically has to be overhauled during a shop visit to restore safety standards and performance. During state of the art shop visits, engines are repaired based on EGT-specifications. To further enhance the maintenance, tailored repairs for each jet engine based on engine history and operation conditions are necessary to take TSFC into account. To ensure such an effective maintenance, the aerodynamic behavior of deteriorated and repaired airfoils is the key factors. Therefore, geometric properties with high influence on aerodynamic performance have to be known. For blisks (BLade-Integrated-diSK) the approach of tailored maintenance will be even more complicated because the airfoil arrangement cannot be changed or individual airfoils cannot be replaced. Thus, the effects of coupled misshaped airfoils have a high significance. This study will present a Design of Experiments (DoE) for circumferential coupled HPC-airfoils to identify the geometric properties which lead to a reduction of performance. To focus on geometric variations, quasi3D (Q3D) simulations are taken out. Based on a sensitivity analysis, the thickness related parameters, the stagger angle as well as the max. profile camber thickness are identified as the most important parameters which are influencing adjacent airfoils and reduce the aerodynamic performance.

Impacts of foreign objects can cause cracks and dents in airfoils, especially in the leading edge. The regeneration of high-pressure compressors blisks with current repair methods is often restricted to a local blending of these edges. This can cause significant changes in the airfoils’ geometrical properties, which in turn influence their aerodynamic and aeroelastic characteristics. Changes at the leading edge have a particularly strong influence on the airfoils’ aerodynamic properties. In order to be able to make an informed decision about if and how a repair should be performed, consequences have to be predicted in advance. To investigate their influence on the aerodynamic and aeroelastic behavior, typical blend repairs are applied to the geometry of a blisk in a 1.5-stage research axial compressor [1], which are representative in shape and size. Blisks (Blade-Integrated-diSK) are function integrated components, which are expected to have a high life span due to significant costs in design and production. Similar modifications are implemented at different radial heights of the blades, in order to investigate the influence of location and penetration depth of blend repairs. It is assured that only the blend repair region is modified while the rest of the blade stays in the original shape. Thus, a realistic change of the geometry is given. The numerical study presented here deals with the influence of geometric imperfections, blend repairs in particular, on the aerodynamic and aeroelastic behavior of the high pressure compressors blisks. Results show that blend repairs have an influence on the local pressure distribution as well as on the local flow turning. Even though the leading edge is reshaped during repair, performance degradation can be observed. Furthermore, the working range of the compressor stage is influenced by the blend-repairs, which is of great importance for safe operation. Finally, the local changes in aerodynamics and blade deformation influence the aeroelastic behavior. This influence depends on the investigated mode shape and the location of the modification. The closer the modification is located towards the tip, the more pronounced are the shifts in aerodynamic damping and aerodynamic stiffness. Low torsional mode shapes display the highest sensitivity to the modifications.

The performance of a compressor decreases during a flight operation interval due to deterioration effects. This results in an increase of specific fuel consumption (SFC) and exhaust gas temperature (EGT), both having an influence on the On-Wing-Time of a jet engine and hence the maintenance intervals. A significant type of deterioration is erosion leading to misshaped airfoils such as thinner or thicker leading- and trailing edges, thinner airfoils as well as a reduction in chord length and an increase in tip clearance. Another deterioration effect are changes in stagger angles. The parameters mentioned above vary between each airfoil in a compressor stage. For a BLISK (BLade-Integrated-diSK) the arrangement of blades in each stage cannot be changed during maintenance, so these effects will grow stronger until the part reaches its life limit. Therefore, the objective of this study is to analyze the effect of circumferential different types of misshaped airfoils on the performance of a compressor stage. A conventional high pressure compressor stage is to be simulated with maximum leading edge thickness in combination with maximum and minimum stagger angle using realistic boundary conditions. In general, the coupled effects due to detoration lead to mistuned throttle lines and a different operation range. Further the geometry of a BLISK being at its life limit is digitalized by a structured light 3D scanner in conjunction with a photogrammetry system to examine the state of detoration and evaluate if the results of a conventional compressor stage can be applied to the new BLISK design. Thus, this study gives a preview on the changes of the expected BLISK performance caused by coupled deterioration effects.