Eksenel kompresörlerde oluk genişliği ve en-boy oranının aerodinamik performansa etkisinin HAD ile analizi

Abstract

Türbomakinelerde verimi ve operasyonel kararlılığı sınırlandıran en temel sorunlardan biri, rotor kanat ucu bölgesinde meydana gelen ikincil akışlar ve buna bağlı olarak ortaya çıkan stall eğilimidir. Kanat ucu boşluğu, mekanik toleransları güvence altına almak amacıyla gerekli olmakla birlikte, bu boşlukta oluşan basınç farkı nedeniyle gelişen sızıntı akışı ve vorteksler kompresör performansında ciddi kayıplara yol açmaktadır. Literatürde bu tip kayıpları azaltmaya yönelik çeşitli pasif kontrol yöntemleri önerilmiş olup, özellikle çevresel oluk uygulamaları son dönemde öne çıkmıştır. Ancak oluk geometrisinin, özellikle genişlik ve en boy oranı gibi parametrelerinin kompresör performansı üzerindeki etkileri hakkında sistematik ve doğrulanmış sayısal ya da deneysel çalışmalar yetersizdir. Bu tez kapsamında, kanat ucu bölgesinde sızıntı akışlarını kontrol etmek ve stall marjını iyileştirmek amacıyla dokuz farklı oluk konfigürasyonu oluşturularak parametrik HAD analizleri gerçekleştirilmiştir. Oluk genişliği yüzde 3, yüzde 6 ve yüzde 9; en boy oranı ise 1, 1.5 ve 2 değerlerine ayarlanarak farklı konfigürasyonlar tasarlanmış ve bunlar oluksuz referans geometriyle karşılaştırılmıştır. Sayısal çalışmalar için NASA tarafından deneysel verileri mevcut olan Rotor 37 eksenel kompresör modeli referans alınmıştır. Deneysel veriler özgün bir test düzeneğinde elde edilmiş olup, HAD modellemesinin doğruluğunu ve çözüm ağı bağımsızlığını sağlamak amacıyla temel kıyaslama aracı olarak kullanılmıştır. Analizlerde SST k-ω türbülans modeli tercih edilmiş ve yaklaşık 1.5 milyon elemandan oluşan çözüm ağı ile kritik bölgelerde düşük y+ değerleri hedeflenmiştir. Parametrik çalışmalar sonucunda, özellikle yüzde 3 ve yüzde 6 genişlikte ve düşük ya da orta en boy oranına sahip oluk konfigürasyonlarının stall marjını yüzde 15 ila 19 aralığında artırdığı ve verimde ise yüzde 0.1'den daha düşük seviyede kayıplara yol açtığı belirlenmiştir. Buna karşılık, yüzde 9 genişlikteki veya yüksek en boy oranına sahip oluk konfigürasyonlarında ana akışın oluk bölgesinde bozulduğu, uç bölgede düşük enerjili alanların ve vortex yapılarının büyüdüğü; buna bağlı olarak hem stall margininde hem de verimde belirgin düşüşler yaşandığı tespit edilmiştir. Akış görselleştirmeleri, olukların uç bölgesinde oluşan sızıntı akışlarını ve vorteks yapısını baskılayarak daha kararlı bir akış rejimi oluşturduğunu göstermiştir. Q-kriteri analizlerinde de, oluklu ve oluksuz geometrilerdeki vorteks yapılarının boyut ve şiddetindeki değişimler karşılaştırılmıştır. Sonuçlar, optimum oluk tasarımının pasif bir kontrol yöntemi olarak kompresörlerin operasyonel sınırlarını genişletebileceğini ve çok düşük verim kaybı ile daha kararlı çalışma aralığı sağlayabileceğini göstermiştir. Bu çalışma, ileride oluk lokasyonu, genişliği, derinliği ve sayısı gibi parametrelerin optimizasyonu ile ileri düzey sayısal ve deneysel yöntemlerle araştırılmasının önünü açacak bir temel sağlamıştır

One of the primary limitations that reduce the aerodynamic performance and operational stability of axial compressors is the development of secondary flows and tip leakage vortices in the rotor blade tip region. The unavoidable clearance between the rotor blade tips and the casing wall is necessary for mechanical and thermal tolerances. However, this clearance gives rise to a pressure differential across the blade tip, resulting in leakage flow from the pressure side to the suction side. The resulting tip leakage vortex interacts destructively with the main flow, contributing to increased energy losses, flow separation, and aerodynamic instabilities such as rotating stall and surge. Stall is a critical flow instability in axial compressors and is often triggered by boundary layer separation or excessive blade loading. Once the flow begins to separate, a cascade of performance losses occurs, including reduced pressure ratio, loss of efficiency, and high-amplitude pressure fluctuations. If the instability progresses further, it may evolve into surge, where the flow reverses direction, potentially causing severe damage to the machine. Thus, extending the stall margin and enhancing compressor stability are fundamental challenges in turbomachinery design. To address these challenges, numerous active and passive flow control techniques have been proposed. Active control methods—such as blowing/suction systems, synthetic jets, and plasma actuators—can directly influence the flow, but they often require additional energy input and are complex to implement. In contrast, passive techniques, which alter the flow path or geometry without requiring extra power, offer a cost-effective and reliable alternative. Among these, circumferential casing grooves have received considerable attention due to their simplicity and potential to delay stall without compromising efficiency. Circumferential grooves modify the near-wall flow structure, particularly in the tip region, by introducing cavities that interact with the tip leakage flow. Depending on their geometry and placement, they can suppress vortex formation or alter vortex trajectory, thereby reducing the negative impact of leakage flow on compressor performance. Although several studies in the literature have examined the effectiveness of grooves in general, detailed parametric studies focusing on groove width and aspect ratio (width-to-height) have been limited. This thesis presents a systematic computational fluid dynamics (CFD) analysis to evaluate the aerodynamic impact of casing groove geometry on axial compressor performance. A total of nine groove configurations were analyzed, resulting from combinations of three groove widths—3%, 6%, and 9% of chord length—and three width-to-height ratios—1.0, 1.5, and 2.0. The baseline geometry used for the analyses was the well-documented NASA Rotor 37, a transonic single-stage axial compressor rotor with abundant experimental data available for validation purposes. The simulations were conducted using the ANSYS CFX CFD software, with high-resolution structured meshes created in ANSYS TurboGrid. Approximately 1.75 million cells were used in each simulation, with particular focus on achieving low wall y⁺ values (<3) near the blade surfaces and casing to accurately capture the boundary layer and tip leakage interactions. The SST k-ω turbulence model was employed due to its well-established ability to handle flow separation and wall-bounded flows in turbomachinery applications. The CFD model was rigorously validated by comparing the baseline (grooveless) simulation results with experimental measurements of Rotor 37, including pressure ratio, mass flow rate, and adiabatic efficiency. The close agreement between simulation and test data confirmed the mesh independence and robustness of the numerical model. The parametric study revealed several important findings. Groove configurations with 3% and 6% chord width and aspect ratios of 1.0 and one of 1.5 demonstrated a substantial stall margin improvement of 3–19%, with minimal efficiency loss (less than 0.1%). These configurations succeeded in altering the leakage flow in a favorable manner, suppressing the formation and propagation of tip leakage vortices. Flow visualizations, including streamline plots and pressure contours, confirmed that the grooves helped reattach the separated flow and reduce the size and strength of vortices in the tip region. In contrast, groove configurations with a 9% width or an aspect ratio of 2.0 were detrimental to performance. In these cases, the grooves acted as flow disturbance sources, disrupting the main flow and allowing the growth of low-energy flow zones and strong vortex cores. These effects not only reduced the stall margin but also degraded pressure ratio and overall efficiency. The simulations showed that overly large or deep grooves could cause flow separation within the cavity and obstruct smooth passage of the main flow. The Q-criterion, a vortex identification method based on the second invariant of the velocity gradient tensor, was used to quantitatively assess the strength and distribution of vortices in each configuration. Results from Q-criterion plots showed a clear reduction in vortex intensity and size in the optimal groove configurations compared to both the grooveless baseline and the larger groove cases. This confirmed that well-designed grooves could effectively reduce the intensity of tip vortices and contribute to a more stable aerodynamic environment. These findings suggest that optimal groove geometry can provide meaningful improvements in axial compressor stability with negligible performance penalties. Specifically, the ability to extend the stall margin without introducing significant aerodynamic losses is of great value in aerospace and energy systems, where safety and operational robustness are paramount. Additionally, the passive nature of grooves means they require no power input and have no moving parts, making them ideal for applications where simplicity and reliability are essential. The thesis also establishes a robust simulation methodology for future research. The integration of validated CFD modeling with parametric geometry studies provides a solid framework for expanding the research into other groove parameters, including axial location, groove shape (e.g., rectangular, rounded, or inclined), depth variations, and multi-groove configurations. Furthermore, future work could leverage optimization algorithms such as genetic algorithms (e.g., NSGA-II), surrogate modeling (e.g., RBFNN), or machine learning techniques to explore the design space more efficiently and identify globally optimal groove configurations. In addition to geometric optimization, future investigations may also consider unsteady flow effects, particularly stall inception and vortex dynamics, through unsteady RANS (URANS) or Large Eddy Simulation (LES) techniques. These higher-fidelity approaches can provide deeper insights into time-resolved vortex behavior and tip leakage interactions, which are beyond the reach of steady-state simulations. In conclusion, this study confirms the aerodynamic potential of circumferential casing grooves when designed with careful consideration of key geometric parameters. It contributes valuable knowledge to the body of work on passive flow control in axial compressors and sets the stage for future developments aimed at improving the efficiency, stability, and reliability of turbomachinery system