Rotor Blade Tip Research Articles

A loss reduction design strategy for a ducted fan under boundary layer ingestion condition based on a lip bump

A distributed fan propulsion system can potentially achieve a 10%-20% fuel burn reduction compared with traditional podded engines. However, the ducted fan is continuously working under inflow distortion conditions caused by boundary layer ingestion, thus resulting in an increase of aerodynamic loss and decreasing the aircraft overall aerodynamic benefit achievement. To control the loss in the boundary layer ingestion ducted fan, a loss reduction design strategy based on the lip-installed bump is developed. The computational results show that the lip-installed bump can effectively suppress the intensity of the distorted inflow with a 9.7% reduction at the rotor inlet. The reason is that a high-pressure area will be presented on the mount of bump and a low-pressure on both sides of the bump compression surface which pushes the low-momentum flow in the boundary layer away from lip. In this way, the bump energizes the low-momentum flow on the downer wall of the lip and reduces the local boundary layer momentum loss thickness. In the meanwhile, as the distortion intensity is decreased near the downer wall at rotor inlet, the local blade load is also reduced, especially near the blade tip. Further, the decrease of the blade load near the tip leads to a reduction of the normal pressure difference between the pressure and suction surfaces which can eliminate the interaction of the main flow and the tip leakage flow, resulting in a 6% decrease of flow loss in the rotor blade tip region. In comparison to the baseline ducted fan, the fan with a lip-installed bump has achieved a 1.3% enhancement of adiabatic efficiency at the re-design point under inflow distortion condition, without a sacrifice of total pressure ratio.

Effect of Change in Material Properties of the Abradable Coating on the Wear Behavior of It—Microstructure Model-Based Analysis Approach

In aerospace applications, engine parts, especially those around the rotor blade tips, are coated with an abradable seal, a specific material layer. Its design produces a tighter seal without harming the blades by allowing it to wear down or “abrade” somewhat when the blade tips come into contact. In turbines and compressors, this reduces gas leakage between high- and low-pressure zones, increasing engine efficiency. Abradable seals are crucial to contemporary jet engines because they enhance performance and lower fuel consumption. The materials selected for these seals are designed to balance durability and abrasion resistance under high temperatures and speeds. Metal matrix, oxide particles, and porosity are the three most prevalent phases. An ideal mix of characteristics, such as hardness and erosion resistance, determines how effective a seal is, and this is accomplished by keeping the right proportions of elements in place throughout production. The primary objective of this research is to optimize abradability by utilizing various FEM tools to simulate the rub rig test and modify testing parameters, including Young’s modulus, yield stress, and tangent modulus, to analyze their impact on the wear behavior of the abradable seal and blade. Two microstructure models (CoNiCrAlY–BN–polyester coating) were found to perform optimally at porosity levels of 56% and 46%, corresponding to hardness values of 48 HR15Y and 71 HR15Y, respectively. Changing factors like yield stress and tangent modulus makes the seal more abrasive while keeping its hardness, porosity, and Young’s modulus the same. Furthermore, altering the Young’s modulus of the shroud material achieves optimal abradability when tangent modulus and yield stress remain constant. These findings provide valuable insights for improving material performance in engineering applications. To improve abradability and forecast characteristics, this procedure entails evaluating the effects of every single parameter setting, culminating in the creation of the best abradable materials. This modeling technique seems to provide reliable findings, providing a solid basis for coating design in the future.

Blade tip clearance measurement method based on blade tip timing without the once per revolution sensor

Abstract The rotor blade tip clearance (BTC) of an aeroengine is a critical parameter that significantly impacts its performance and safety. Accurately determining dynamic BTC has been a focal point of research in experimental testing. Among the effective measurement methods for dynamic BTC, the laser triangulation method based on blade tip timing (BTT) stands out. However, this method typically relies on an once-per-revolution (OPR) sensor for rotor speed and blade serial number information during measurement, leading to increased installation and maintenance costs as well as additional measurement uncertainties. In response to this challenge, this paper presents two signal processing method that eliminate the need for an OPR sensor. The first method, referred to as ‘none OPR method 1,’ closely follows the measurement principles of the traditional method but employs a skewed dual light probe (SDLP) signal to extract rotational speed information, thus obviating the need for an OPR sensor. The second method, ‘none OPR method 2,’ introduces a novel data processing approach that enables BTC determination without any source of rotor speed information. Both methods utilize dynamic and static BTC matching techniques to identify blade serial numbers and achieve BTC measurement without an OPR sensor on a simulated rotor experimental bench. Comparative analysis with traditional method BTC measurements reveals that, under stable rotor speeds, all methods achieve similar levels of accuracy and repeatability errors within 0.025 mm. However, when dealing with variable speed conditions, the traditional method’s accuracy is significantly affected. None OPR method 1 mitigates the impact of rotor speed changes to some extent, while none OPR method 2 remains unaffected by such changes. Importantly, both none OPR methods demonstrate the capability to measure blade-by- blade tip clearances akin to the traditional method.

The Influence of Bleed Position on the Stability Expansion Effect of Self-Circulating Casing Treatment

The self-circulating casing treatment can effectively expand the stable working range of the compressor, with little impact on its efficiency. With a single-stage transonic axial flow compressor NASA (National Aeronautics and Space Administration) Stage 35 as the research object, a multi-channel unsteady numerical calculation method was used here to design three types of self-circulating casing treatment structures: 20% Ca (axial chord length of the rotor blade tip), 60% Ca, and 178% Ca (at this time, the bleed position is at the stator channel casing) from the leading edge of the blade tip. The effects of these three bleed positions on the self-circulating stability expansion effect and compressor performance were studied separately. The calculation results indicate that the further the bleed position is from the leading edge of the blade tip, the weaker the expansion ability of the self-circulating casing treatment, and the greater the negative impact on the peak efficiency and design point efficiency of the compressor. This is because the air inlet of the self-circulating casing with an air intake position of 20% Ca is located directly above the core area of the rotor blade top blockage, which can more effectively extract low-energy fluid from the blockage area. Compared to the other two bleed positions, it has the greatest inhibitory effect on the leakage vortex in the rotor blade tip gap and has the strongest ability to improve the blockage at the rotor blade tip. Therefore, 20% Ca from the leading edge of the blade tip has the strongest stability expansion ability, achieving a stall margin improvement of 11.28%.

Influence of tip air injection on the unsteady blade force under circumferential pressure distortion

In order to further understand the impact of tip air injection on compressor stability under circumferential pressure distortion, a series of experimental studies on tip air injection under uniform and circumferential distortion were conducted on a low-speed single rotor axial flow compressor. The unsteady measurements were carried out by use of microsensors and strain gauges bonded on the rotor blade surfaces, and a collection of dynamic pressure sensors installed on the casing wall. Results indicate that with the increase in the injected momentum ratio, the load at the leading edge near the rotor blade tip obviously decreases, thereby delaying the stall. By comparing the pressure distribution on the rotor blade surface under different injected momentum ratios, tip air injection can effectively suppress flow separation at the blade tip to enhance flow capability. In addition, it can also weaken the fluctuation of tip leakage flow and reduce the dynamic stress similar to natural vibration near the blade root both in stationary and rotating coordinate systems. Under the condition of inlet circumferential distortion, tip air injection can significantly reduce the load at the distorted region and suppress the low-frequency disturbances measured on the rotor blade surface caused by inlet distortion. In addition, the dynamic stress near the blade root can also be effectively reduced by tip air injection. It implies that tip air injection can both extend the stall margin and reduce blade vibration when suffering from circumferential distortion; thus, the compressor can safely operate.

Aerothermal Performance Robustness and Reliability Analysis of Turbine Blade Squealer Tip With Film Cooling

Abstract The uncertainty quantification in the turbine components' aerodynamic and heat transfer performances is widely considered to be the most challenging topic due to its intricate and nonlinear characteristics. This paper first proposes an efficient uncertainty quantification method based on an original parallel framework combining Polynomial Chaos Expansions (PCE) with two forms (stochastic response surface-based and Galerkin projection-based) and the Universal Kriging method. The rigorous mathematical tests are performed to verify the reliability and computational efficiency of the proposed method, and the results support that this method can dramatically reduce computational samples compared to the conventional PCE method while maintaining computational accuracy. Then, the genetic algorithm was introduced to establish an efficient uncertainty quantification framework, and it is applied to the aerothermal performance robustness investigation of the GE-E3 rotor blade tip with and without film cooling. Based on the findings of uncertainty quantification, the injection of cooling air drastically enhances the unstable tendency of the flow and thermal fields, resulting in the actual aerothermal performance of the squealer tip being much lower than that predicted by deterministic calculations. The setting of the film cooling, although effective in reducing the heat flux around the cooling holes, also induces more chaotic flow and thermal fields, leading to sharp heat flux fluctuations around the cooling holes. Finally, our novel reliability analysis algorithm, rooted in the quantification of uncertainty, corroborates the assertion that the introduction of coolant gas, while extending the operational longevity of turbine blades, confers only marginal improvements in the mitigation of lifespan variability. The comprehensive lifespan assessment elucidates that the mean operational longevity of the conventional squealer tip design stands at an estimated 16,169.44 h, accompanied by a standard deviation of 2,750.31 h. In stark contrast, the mean operational longevity of the squealer tip integrated with film cooling measures a significantly enhanced 17,035.17 h, exhibiting a standard deviation of 2,492.73 h. Consequently, the operational lifespan of the conventional squealer tip experiences a decrement of 10.17% in comparison to the anticipated mean lifespan, whereas the reduction for the film-cooled squealer tip registers at 5.36%.

Experiment study on particle pulverization-assisted convective air drying of sewage sludge

Sewage sludge drying is a pivotal step in sludge reduction, harmlessness, and resource utilization. Pulverizing sewage sludge can significantly increase the surface area of sludge particles available for contact with the air medium, consequently strengthening the convective drying process of the sludge. In this paper, the adopted deep drying system mainly combined sewage sludge pulverization and sewage sludge drying functions. Parametric studies were carried out to investigate the effects on the drying process. The moisture content of the sewage sludge decreased from 50.3% to 26.8% when the linear speed of the rotating blade tip of the pulverizing dryer and induced draft fan reached 28.3 m·s−1, 62.8 m−1, and air velocity reached 9.5 m·s−1, respectively. Furthermore, this work also preliminarily reveals the mechanism of the pulverizing dryer. The pores on the sludge particle surfaces are destroyed during the drying process. As the linear speed of the rotating blade tip in the pulverizing dryer increased, the hydrophilicity and the average particle size of the sewage sludge particles decreased, with the particle size representing the highest volume fraction being reduced by 62% in the particle size distribution. The pulverizing dryer not only breaks up the sludge to increase the surface area of the sludge in contact with the air but may also have a centrifugal effect. As the intensity of the pulverizing dryer increases, Nuclear Magnetic Resonance (NMR) measurements show a leftward shift in the peak of the weakly bound water in the sludge and a noticeable shortening of the transverse relaxation time. Overall, the results suggested that pulverizing is a suitable assistive method to improve the drying effect of the sewage sludge.

Fluctuation of the unsteady force on the rotor blade surface under circumferential distortion in axial flow compressor

To deeply understand the negative impact of circumferential inlet distortion on the internal flow of the compressor, by using the unsteady force testing technique of a rotor blade surface under a rotating coordinate system, the pressure fluctuations on the rotor blade surface are successfully captured when the rotor blade rotates through the distortion region. Results show that the pressure on the suction surface increases and the inlet angle of attack decreases before the rotor blade enters the distorted region. The pressure difference between the pressure and suction surface is obviously enhanced, thereby sharply increasing the blade load and intensifying the flow separation, which is easy to induce instability. When the rotor blade rotates out of the distortion region, the pressure on the suction surface is still low; thus, the inlet angle of attack in this position is larger than before entering the distorted region. It also shows that the outlet of the distorted region is prone to trigger stall. In addition, the dynamic spectrum characteristics of unsteady forces on the blade surface demonstrate that the energy of the rotor frequency and its harmonics increases significantly and the energy of low frequency disturbance is enhanced when the rotor blade rotates through the distorted region. As a result, the vibration is more obvious in the distorted region, especially the energy of natural vibration frequency of the rotor blade is enhanced. When the compressor stalls, the vibration at the rotor blade root is intensified, which is significantly stronger than at the rotor blade tip. It provides support for evaluating the influence of stall and surge on the lifecycle of rotor blade when suffering from circumferential inlet distortion.

Distance Measurement of Contra-Rotating Rotor Blades with Ultrasonic Transducers.

Coaxial rotor helicopters have great potential in civilian and commercial uses, with many advantages, but challenges remain in the accurate measurement of rotor blades' distance to prevent blade collision. In this paper, a blade tip distance measurement method based on ultrasonic measurement window and phase triggering is proposed, and the triggering time of the transmitter is studied. Due to the complexity of the measured signal, bandpass filtering and a time-of-flight (TOF) estimation based on the power density of the received signal are utilised. The method is tested on an experimental test platform with a pair of 200 kHz ultrasonic transducers. The experimental results show that the maximum ranging error is less than 1.0% for the blade tip distance in a range of 100-1000 mm. Compared with the amplitude threshold method, the proposed TOF estimation method works well on the received signal with a low SNR and improves the ranging accuracy by about 5 mm when the blade tip distance is larger than 500 mm. This study provides a good reference for the accurate measurement of rotor blade tip distance, and gives a solution for ranging high-speed rotating objects.

Experimental Investigation of Tip Design Effects on the Unsteady Aerodynamics and Heat Transfer of a High-Speed Turbine

Abstract While modern engine manufacturers devote significant efforts to the development of reliable and efficient machines, the introduction of novel, optimized components in the hot gas path represents a risky opportunity. Accurate experimental and numerical data are critical to assess the impact of new technologies on the harsh engine environment. The present study addresses the impact of a selection of high-performance rotor blade tips on the aerodynamic and heat flux field of a high-pressure turbine (HPT) stage. A combined numerical and experimental approach is employed to characterize the interaction of the tip leakage flow with the rotor secondary flows and the casing heat transfer mechanisms for each individual tip geometry. The turbine stage is tested at engine-scaled conditions in the rotating turbine facility of the von Karman Institute. In the present study, the turbine rotor is operated in rainbow configuration to allow the simultaneous testing of multiple blade tip geometries. Reynolds-averaged Navier–Stokes (RANS) simulations are employed to predict the aerodynamic and thermal fields of the individual profiles using test-calibrated boundary conditions. Isothermal steady computations are performed at different wall temperatures to compute the adiabatic wall temperature and the heat transfer convective coefficient. Low-order models are used to represent the over-tip thermal field and the driving heat transfer mechanisms. The time-resolved outlet flow is characterized using a vortex tracking technique and high-frequency aerodynamic measurements to identify the rotor secondary flow structures.

Design Optimization of Blade Tip in Subsonic and Transonic Turbine Stages—Part I: Stage Design and Preliminary Tip Optimization

Abstract Rotor blade tip has significant influence on turbine stage aerodynamics and heat transfer. Most previous efforts have been based on low-speed cascade settings. However, more recent research on transonic blade tips exhibits distinctive flow features with qualitatively different performance sensitivities. These prompt two key issues of interest on the related flow conditioning. First, the contrast between a low-speed flow and a transonic regime highlights the less studied high-subsonic flow regime, closely relevant to many realistic turbine designs. Second, the relative casing movement and upstream inflow conditions, known to have non-negligible effects, indicate the need to examine a rotor blade tip in a realistic stator–rotor stage environment, which is also lacking. To elaborate the Mach number effect in the flow regimes of practical interest, we aim to examine a high subsonic stage in a direct and consistent comparison with a transonic one. To this end, a high subsonic stage (exit Mach number of 0.7) and a transonic (exit Mach number of 1.1) are designed at the same Reynolds number with a three-dimensional parameterization and meshing system. The tip squealer height is used as a representative parameter to investigate the sensitivity of the stage aerothermal performance. The multi-objective optimization using the Kriging surrogated model is employed to identify the Pareto fronts for the stage efficiency and the heat transfer. The comparison of the optimized results between these two stages shows distinctively different trends in the performance variation with the squealer height. The efficiency of the subsonic stage increases with the squealer height reaching a plateau. In contrast, the efficiency in the transonic stage first increases and then drops to the level comparable to that of a flat tip. Significantly, the present results indicate, for the first time, that the squealer tip in a transonic stage may not be as effective as in a subsonic stage. On the other hand, for heat transfer, sensitivity variations are more complex. The overall heat load and the local nonuniformity lead to qualitatively different sensitivities with the squealer height, as well as completely incomparable Pareto fronts. These observed heat transfer sensitivities raise the question on how to effectively conduct a combined aerodynamic and heat transfer performance design optimization. The authors subsequently resort further aerothermal physics analyses described in a companion paper as Part II of the two-part article. In Part II, the physical interpretation of the contrasting aero-efficiency sensitivities for the two stages, as well as a physical understanding leveraged selection of the objective function for such combined blade tip aerothermal optimization will be presented.

Effect of rotor blade position on Vertical Axis Wind Turbine performance

In this paper a numerical study is presented with the aim of evaluating the performance output of three Vertical Axis Wind Turbine (VAWT) configurations. Here using Computational Fluid Dynamics (CFD) a two-dimensional Multi Reference Frame (MRF) approach has been used to perform steady state simulations. For this purpose an inlet velocity of 4m/s has been used along with rotor blade tip speed ratios (λ) in the range of 0 to 0.6. The effects of varying rotor blade position on the global and local flow fields have been quantified. Furthermore, the influence of rotor blade position on performance has been computed. It has been found that within a typical rotor blade passage, maximum torque is obtained at a unique rotor angle. It can be seen that the torque output of the turbine decreases with an increase in rotor tip speed ratio (λ) for all rotor blade positions. From the data obtained it can be concluded that the VAWT power curve characteristics vary with relative rotor blade position.

Controlling the flow loss of a supersonic compressor rotor using the blade slotting method

The flow loss in the blade passage of a supersonic compressor rotor mainly comes from the boundary layers on the blade surface and end wall, the shock wave, the shock wave/boundary layer interaction, and tip leakage flow; the instability is mainly caused by the shock wave near the rotor blade tip exiting the blade passage. This paper adopts an internal slot in the blade, with the inlet of the slot located at the leading edge of the blade and the outlet located on the suction surface of the blade, by using the momentum of the incoming flow to form a high-velocity jet to control the flow loss and improve the stall margin of the supersonic rotor. The mechanism of reducing flow loss by a slotting jet was studied, and a genetic algorithm optimization platform was further used for the coupled optimization design of the slot and blade. The numerical calculation results showed that the slotting jet can effectively suppress the development of the boundary layer on the suction surface while reducing the intensity of the shock wave, thereby reducing the loss of the boundary layer and shock wave, significantly improving the peak efficiency of the rotor, and increasing the mass flow rate at the peak efficiency point. The slotting jet can cause the shock wave in the passage to move downstream, thereby improving the stall margin of the rotor. Due to the strong shock wave in the blade passage near the blade tip, the slot outlet should be near and upstream of the shock wave; the shock wave in the middle and root regions of the blade is weaker, and the slot outlet should be located downstream of the shock wave.

Effect of rotor blade tip clearance on aerodynamic performance of high-speed multi-stage axial compressor

Effect of rotor blade tip clearance on aerodynamic performance of high-speed multi-stage axial compressor

Optimization of Rotor Tip Cavity Shapes for Mitigating Aerodynamic Tip Leakage Losses

Rotor blade tip leakage loss influences turbine stage efficiency. A parametric optimization method based on five design variables combined with Non-Linear Programming by Quadratic Lagrangian to optimize the turbine efficiency estimated from computational fluid dynamics analysis has been used to obtain estimates of the optimal cavity depth and split rib location within the tip cavity, which appears to improve the stage efficiency by 1.5%. Besides the design point, the influence of the optimized tip cavity shape impacts the performance of the rotor at different rotational speeds and pressure ratios. In this study, by comparing different tip cavity shapes, it is shown that a dual-cavity shape on the rotor blade tip with a stream-direction split rib can effectively control the air flow bending angle inside the tip clearance and improve the aerodynamic performance. The compressible unsteady Reynolds-averaged Navier–Stokes model is used to verify the adaptability of the optimized tip cavity shape to periodic wake disturbances, and the results demonstrate that the time-averaged estimate of the efficiency can be improved by 1.3% compared to that of the flat rotor tip.

Partial tip air injection to extend the stall margin in an axial flow compressor with circumferential inlet distortion

As a typical non-axisymmetric inflow, circumferential pressure inlet distortion was experimentally investigated in an axial flow compressor. Based on the obvious decreasing trend of stall margin with the enhancement of distorted intensity, the stall mechanism under the condition of maximum distorted intensity, namely, 24.57 %, was studied in detailed. The casing static pressure over the rotor blade tip along the chord and circumferential direction were measured using a collection of dynamic transducers. Results show that under the circumferential inlet distortion condition, the pressure disturbance is always generated downstream from the distorted region along the rotational direction, and circumferentially propagates until it disappears in the upstream of the distorted region. With the continuous throttling, if the pressure disturbance that has not completely disappeared in the undistorted region enters the distorted region again, its energy will be enhanced, finally induces the stall. Using the above understanding as a foundation to apply partial tip air injection to suppress the disturbance caused by the circumferential inlet distortion, and resist its adverse effect on stall margin to ensure the safe operation of the compressor. It demonstrated that the tip air injection applied in the downstream of the distorted region can obtained the largest stall margin on the premise of the same injected momentum. For the strongest distortion case (DC (60) equals to 24.57 %), the injectors fixed downstream the distorted region can reduce the number and scale of the spike disturbances, suppress the local tip leakage vortex, and decrease the amplitude of the pressure disturbance energy, thus a maximum stall margin improvement of 3.18 % is obtained. However, tip air injection in the undistorted region cannot contribute the stall margin improvement that it leads to stronger tip leakage vortex and earlier stall inception. This study provides a guideline to design the non-axisymmetric tip air injection for circumferential inlet distortion with a help of injected mass flow and energy reduction.

Numerical analysis of the influence of hull-modulated inflow on unsteady force fluctuations and vortex dynamics of pump-jet propulsor

The influence of the hull-modulated inflow on the propulsion performance of the propeller is related to the matching design of the propeller–hull system. In the present study, considering the working conditions of the pump-jet propulsor in uniform inflow and two types of hull-modulated inflow, based on improved delay detached eddy simulation, the influence of hull-modulated inflow on unsteady force fluctuations and vortex dynamics of pump-jet propulsor under design conditions is carried out. The results show that the hull-modulated inflow increases the propulsion efficiency of the pump-jet propulsor to varying degrees within the range of the calculated advance coefficient and has a significant influence on the frequency characteristics of the unsteady force spectra characteristics of each component of the pump-jet propulsor. It also shows changes in the magnitude characteristics, that is, the energy transfer process of an individual rotor blade from the stator blade passing frequency to other harmonics of the shaft rotation frequency, and the thrust spectrum of an individual stator blade presents broad-spectrum characteristics in the high-frequency range. Furthermore, the application of hull-modulated inflow directly affects the shape of the stator shedding vortex, causing some of the stator blade shedding vortices to separate early and aggravating its short-wave instability. More secondary vortices are induced to accelerate the instability of the rotor blade tip clearance vortex. The energy transfer mechanism from the rotor blade passing frequency and its harmonics to the broadband spectra appears in the wake field of the pump-jet propulsor.

Effect of wire mesh casing treatment on axial compressor performance and stability

In this paper, a kind of Wire Mesh Casing Treatment (WMCT) is proposed to improve the stable operating range of the compressor. In contrast to the traditional circumferential groove, as for WMCT, a layer of wire mesh is laid on the surface of the circumferential groove. Parametric studies were conducted on the low-speed axial flow compressor, including the groove width, axial location, and mesh count. The optimum axial location for WMCT is related to its groove width. A higher wire mesh count results in a smaller compressor stall margin improvement. Steady simulations were carried out to study the effect of WMCT on the flow structure of the compressor. The wire mesh in the WMCT has a certain flow resistance, which restricts the flow into and out of the groove. Due to the WMCT, the flow parameter in the tip region of the rotor is less sensitive to changes in the operating conditions of the compressor. The WMCT causes the rotor tip blade loading to shift backward, inhibiting the formation of spill forward of the leakage flow, and thus improving the stability of the compressor. The flow resistance on the groove surface is a new degree-of-freedom for the casing treatment designer.

Numerical study of airfoil used for helicopter rotor tip to investigate the enhancement in the aerodynamic characteristics using active flow control at forward flight conditions

Helicopter forward flight is the most important state of flight due to time spent in this mode during the complete helicopter trip. In this study, a main rotor blade tip airfoil was simulated to obtain the effect of applying active flow control on the tip part of the rotor blade. A blowing excitation jet is applied to the leading edge of the airfoil just before the separation point with small momentum and high frequency to enhance the near wall boundary layer characteristic. This jet is used with different amplitude, frequency and excitation location from a leading edge to study the effect of this parameter. A numerical model is used to solve two different angles of attack and all results are compared with NACA report that used the same helicopter hub and airfoil. Results show good agreement with the experimental results and the chosen excitation parameter gives good enhancement to the performance of the airfoil used in the tip part of the helicopter blade in this mode of flight. Performance represented in lift and drag and lift to drag ratio shows improvement, reaching about 10% increase in lift and 30% decrease in drag, which means more than 50% change in lift to drag ratio. Furthermore, this study gives a good estimation of the best excitation location that is required for real application on the helicopter blade to improve of forward flight performance.

Influence of the chordwise distribution of tip winglets on the stability of a high-load compressor stage

Compressor rotor blade tip winglet technology has been confirmed to be a passive flow control method that is effective in increasing the operating stability of compressors. As revealed by extensive research, the structural parameters of tip winglets are capable of affecting their stability expansion effect, whereas their mechanism and influence law of the compressor stage remains unclear. In this study, the effect exerted by the chordwise distribution of tip winglets on the stability of a high-load compressor stage was investigated. As indicated by the result of this study, the tip winglet exhibiting a proper widest position and an adequate chordwise length can effectively increase the stall margin. To be specific, the widest position was located at 25% chord, and the stall margin was increased by up to 25.39%. A winglet exhibiting a chordwise length of 50% achieved the stability expansion ability equivalent to winglet with a full chordwise length, which can be increased by 22.58%. The tip leakage flow turned out to be weakened as the axial momentum was increased, and the circumferential momentum of the flow that arose from the tip winglet was reduced. The result of this study also suggested that tip winglets can play a certain role in combing the inflow of the stator, such that the blockage of the stator channel can be reduced.