Rotor Tip Gap Research Articles

Effects of the rotor tip gap on the aerodynamic and aeroacoustic performance of a ducted rotor in hover

Ducted rotors are configurations known to outperform their unducted reference baselines when aerodynamic performance is concerned. Aside from aerodynamic benefits in hover, a duct also affects acoustic emissions. One of the most contended design parameters of a duct-rotor assembly is the radial distance between the blade tip and the duct wall, referred to as the “tip gap”. The present study explains how the aerodynamic performance of a ducted-rotor system is affected by the tip-gap distance, taking into account the performance of the rotor and those of the duct's inlet lip and diffuser sections. Separate thrust measurements of the rotor and duct establish that the latter can generate up to half of the total thrust of the assembly. Static wall-pressure measurements along the inner wall of the duct reveal a low pressure suction zone over the duct's inlet lip area. This allows the assembly to generate more thrust than the rotor alone, even though the duct's diffuser section generates a drag component (negative thrust). From the velocity fields it is further shown that the performance-deterioration with an increasing tip gap distance is associated with a contraction of the rotor slipstream in the duct diffuser.

Leading-edge flow prediction in Wells turbine under the influence of tip leakage flows

Monitoring of aerodynamic parameters of the blade helps to enhance the operating capabilities of the Wells turbine. In this study, a few discrete pressure points were used to estimate the pressure patterns by fitting the exact potential-flow solution for flow over a parabola. The sectional aerodynamics was assessed by the leading-edge flow sensing (LEFS) algorithm. A characteristic parameter was employed to deduce leading-edge flow status. Moreover, the influence of rotor speeds and tip gaps on the vortex boundary was discussed in terms of the Lagrangian frame. For the rotor speed larger than 1000 rpm at the incipient stall point, there is no boundary layer separation at the suction side of the blade tip leading edge. At the same rotor speed, a large tip gap reduces the axial shear of the suction flows and the axial stretching of the tip leakage vortex, which helps to enhance the interaction time and strength of the leakage vortex and the suction flows. For the Wells turbine with attached flows, the LEFS algorithm can accurately evaluate the pressure distribution and suction parameters near the leading edge. The LEFS-derived dimensionless parameter can act as an equivalent angle of attack under different tip gaps, providing the possibility for stall warning of Wells turbines.

Inception of Casing Wall Flow Instabilities and Downstream Flow Evolution in Subsonic Axial Compressor Rotors At High Rotor Stagger Angles

Two critical aspects of the casing wall flow structure in a subsonic axial flow compressor are studied in this paper: flow mechanism leading to stalling at different stagger angle configurations; and the effect of tip leakage flows on these flow phenomena. In some of the previous works, casing wall flow instabilities similar to the ones discussed in the present paper were reported, typically for high stagger angle rotor designs. However under what conditions flow tends to take such shapes were not discussed in detail. This work attempts to address this through a detailed numerical study on unswept and forward swept rotors, for different stagger angle settings, and for varied levels of rotor tip gaps. Three candidate rotors (unswept, 20° true swept, 20° axially swept) are studied with three tip clearances (0.0%, 0.7%, and 2.7% of the blade chord) for three stagger angle changes (0°, +3° and +5°). Stagger angle is found to have significant effect on suction surface flow separations. Tip clearance is observed to be critical on the inception of casing wall flow instabilities, especially at higher stagger angles. Hub endwall flow structure and the downstream flow evolution are also discussed in the paper.

Numerical Analysis of the Effects of Different Rotor Tip Gaps in a Radial Turbine Operating at High Pressure Ratios Reaching Choked Flow

To operate, radial turbines used in turbochargers require a minimum tip gap between the rotor blades and the stationary wall casing (shroud). This gap generates leakage flow driven by the pressure difference between the pressure and suction side. The tip leakage flow is largely unturned, which translates into a reduction of the shaft work due to the decrease in the total pressure. This paper investigates the flow through the rotor blade tip gap and the effects on the main flow when the turbine operates at a lower and higher pressure ratio with the presence of supersonic regions at the rotor trailing edge for two rotational speeds using computational fluid dynamics (CFD). The rotor tip gap has been decreased and increased up to 50% of the original tip gap geometry given by the manufacturer. Depending on the operational point, the results reveal that a reduction of 50% of the tip gap can lead to an increase of almost 3% in the efficiency, whereas a rise in 50% in the gap penalty the efficiency up to 3%. Furthermore, a supersonic region appears in the tip gap just when the flow enters through the pressure side, then the flow accelerates, leaving the suction side with a higher relative Mach number, generating a vortex by mixing with the mainstream. The effects of the vortex with the variation of the tip gap on the choked area at the rotor trailing edge presents a more significant change at higher than lower speeds. At a higher speed, the choked region closer to the shroud is due to the high relative inlet flow angle and the effects of the high relative motion of the shroud wall. Furthermore, this relative motion forces the tip leakage vortex to stay closer to the tip suction side, generating a subsonic region, which increases with the tip gap height. The leakage flow at lower and higher rotational speed does not affect the main flow close to the hub. However, close to the shroud, the velocity profile changes, and the generated entropy increases when the flow goes through the tip gap.

Heat transfer characteristics of a high-pressure turbine under combined distorted hot-streak and residual swirl: an unsteady computational study

• Unsteady CFD simulations were performed for a transonic turbine stage to reveal the effects of uniform and swirled flows on hot-streak transport. • Two ideal hot-streak maps, radial and rounded, were evaluated against a realistic distorted hot-streak map. • The residual swirl contributed to the homogenisation of the radial hot-streak and the dispersion of the rounded and distorted hot-streaks through the rotor. • The distorted hot-streak showed the most complex transport behaviour through the stage. • The residual swirl was found to intensify the leakage vortex and increase the heat transfer rates near the tip region. In new generation aero-engines, lean-burn combustors are equipped with swirl injectors in order to reduce pollutant emissions. At the exit of these combustors, the flow is dominated by temperature non-uniformity (hot-streak) and residual swirl. Available research has focused mainly on the isolated effects of the residual swirl only or the hot-streak only on the aerodynamic performance or thermal performance of the high-pressure turbine (HPT). Only few studies investigated the combined swirl and hot-streak on the aerothermal performance of the HPT using either idealized uniform or rounded hot-streak topologies. Realistic hot-streaks generated from lean burn combustors have more complex topologies that involves distortions. The present study investigates the effects of residual swirl and distorted hot-streak simultaneously on a first stage of a HPT. Other hot-streak types, uniform and rounded, have been also investigated with and without swirl to perform comparisons with the distorted hot-streak. Unsteady Reynolds-averaged Navier–Stokes (URANS) computations have been conducted to assess the aerothermal performance of a HPT under the influence of different hot-streaks. Results revealed that all hot-streaks without swirl were almost preserved as transported through the vane and altered as transported through the rotor due to the secondary flows. Under the residual swirl, all hot-streaks were remarkably altered at the vane exit and deformed more at the rotor exit. The uniform hot-streak was homogenised through the rotor and the rounded and distorted hot-streaks were dispersed. The distorted hot-streak showed the most complex transport behaviour through the stage. Results also revealed that the leakage flow through the rotor tip gap generated high heat transfer rates, in particular on the rotor blade suction side and tip surface. The leakage flow formed the known leakage vortex at the tip gap exit, which induced high heat transfer rates. The residual swirl was found to intensify the leakage vortex and consequently higher heat transfer rates are generated compared to the uniform flow condition.

Loss evaluation and aerodynamics investigation of an aggressive intermediate turbine duct under off-design conditions

To maximize the performance of the intermediate turbine duct (ITD) under off-design conditions, the loss generation in a one-half stage turbine was quantified using entropy generation and the global entropy generation rate. The numerical results solved by the unsteady Reynolds-averaged Navier–Strokes equations were first verified. Then, the aerodynamic losses within the high-pressure turbine stage were evaluated by efficiency loss under nine operating conditions composed of three rotor speeds and three rotor tip gaps. Finally, the disturbance modes caused by the upstream wake were captured by the dynamic mode decomposition method. Different from the influence of tip gaps, losses of the high-pressure turbine and the ITD are due to the swirl angle display an opposite trend. Under the influence of the interaction between the tip leakage flow and the shroud flow of the ITD, the viscous dissipation and turbulent dissipation increase with a larger tip gap owing to the dominant counter-rotating vortices and secondary flow occurring near the upstream of the ITD shroud. In addition, a large gap seems to enlarge the swirl component of the inflow angle, especially over 80% passage height, leading to greater dissipation losses in these areas. At the ITD inlet, two pairs of counter-rotating vortices at the shroud and the hub are, respectively, captured by the axial velocity mode. Large tip gaps enhance endwall vortices near the shroud and make the up vortex pairs merge into one pair.

Numerical Analysis of the Effects of Different Rotor Tip Gaps in a Radial Turbine Operating at High Pressure Ratios Reaching Choked Flow

Numerical Analysis of the Effects of Different Rotor Tip Gaps in a Radial Turbine Operating at High Pressure Ratios Reaching Choked Flow

Aerodynamic simulation of a high-pressure compressor stage using the lattice Boltzmann method

PurposeThe present paper aims at evaluating the lattice Boltzmann method (LBM) on a high-subsonic high-pressure compressor stage at nominal regime.Design/methodology/approachThe studied configuration corresponds to the H25 compressor operated in a closed-loop test rig at the von Karman Institute. Several operating points are simulated with LBM for two grids of successive refinements. A detailed analysis is performed on the time-averaged flow predicted by LBM, using a comparison with experimental and existing RANS data.FindingsThe finest grid is found to correctly predict the mean flow across the machine, as well as the influence of the rotor tip gap size. Going beyond time-averaged data, some flow analysis is performed to show the relevance of such a high-fidelity method applied to a compressor configuration. In particular, vortical structures and their evolution with the operating points are clearly highlighted. Spectral analyses finally hint at a proper prediction of tonal and broadband contents by LBM.Originality/valueThe application of LBM to high-speed turbomachinery flows is very recent. This paper validates one of the first LBM simulations of a high-subsonic high-pressure compressor stage.

Influence of Gap Detailing on Calculated Unsteady Non-Adjacent Blade Row Aero-Forcing in a Transonic Compressor Stage

Abstract Resonant or close to resonant forced response excitation of compressor blades limits component life time and can potentially lead to high-cycle fatigue failure if the exciting forces are large and damping is insufficient. When numerically quantifying the forcing function by means of simulations, simplifications are typically made in the analysis to reduce complexity and computational cost. In this paper, we numerically investigate how the blade forcing function is influenced by the rotor tip gap flow and by flow across gaps in the upstream variable inlet guide vane row. Unsteady simulations are made using a test rig geometry where a forcing crossing with an excitation from a non-adjacent blade row had previously been measured. The effects of the gaps on the forcing function for the first torsion mode are presented for both the non-adjacent blade row excitation (changes compared with a case without gaps indicating a 20% reduction) and an adjacent excitation (changes indicating an 80% increase in terms of forcing function amplitude comparing with a case without gaps).

Study on the effects of tooth profile design parameters of rotor to performance of vacuum pump

The vacuum pump usually used traditional curves such as the circular, cycloidal curves and their combinations to construct tooth profile. However, to increase efficiency and design flexibility for the vacuum pump, a novel rotor tooth profile for Roots rotor of vacuum pumps is proposed. Which is named “CEIEC” tooth profile and orderly composed of five significant segments, a circular arc for tooth tip, an epicycloid curve with variable extension, an involute, an enveloped epicycloid curve and a conjugated circular arc for tooth root. A numerical example is presented to evaluate the performance indices for proposed vacuum pump, including the hermeticity coefficients of the rotor mesh gap and tip gap.

The Inner Workings of Axial Casing Grooves in a One and a Half Stage Axial Compressor With a Large Rotor Tip Gap: Changes in Stall Margin and Efficiency

Effects of axial casing grooves (ACGs) on the stall margin and efficiency of a one and a half stage low-speed axial compressor with a large rotor tip gap are investigated in detail. The primary focus of the current paper is to identify the flow mechanisms behind the changes in stall margin and on the efficiency of the compressor stage with a large rotor tip gap. Semicircular axial grooves installed in the rotor's leading edge area are investigated. A large eddy simulation (LES) is applied to calculate the unsteady flow field in a compressor stage with ACGs. The calculated flow fields are first validated with previously reported flow visualizations and stereo particle image velocimetry (SPIV) measurements. An in-depth examination of the calculated flow field indicates that the primary mechanism of the ACG is the prevention of full tip leakage vortex (TLV) formation when the rotor blade passes under the axial grooves periodically. The TLV is formed when the incoming main flow boundary layer collides with the tip clearance flow boundary layer coming from the opposite direction near the casing and rolls up around the rotor tip vortex. When the rotor passes directly under the axial groove, the tip clearance flow boundary layer on the casing moves into the ACGs and no roll-up of the incoming main flow boundary layer can occur. Consequently, the full TLV is not formed periodically as the rotor passes under the open casing of the axial grooves. Axial grooves prevent the formation of the full TLV. This periodic prevention of the full TLV generation is the main mechanism explaining how the ACGs extend the compressor stall margin by reducing the total blockage near the rotor tip area. Flows coming out from the front of the grooves affect the overall performance as it increases the flow incidence near the leading edge and the blade loading with the current ACGs. The primary flow mechanism of the ACGs is periodic prevention of the full TLV formation. Lower efficiency and reduced pressure rise at higher flow rates for the current casing groove configuration are due to additional mixing between the main passage flow and the flow from the grooves. At higher flow rates, blockage generation due to this additional mixing is larger than any removal of the flow blockage by the grooves. Furthermore, stronger double-leakage tip clearance flow is generated with this additional mixing with the ACGs at a higher flow rate than that of the smooth wall.

Engine-Scalable Rotor Casing Convective Heat Flux Evaluation Using Inverse Heat Transfer Methods

In this paper, we propose a two-step methodology to evaluate the convective heat flux along the rotor casing using an engine-scalable approach based on discrete Green's functions . The first step consists in the use of an inverse heat transfer technique to retrieve the heat flux distribution on the shroud inner wall by measuring the temperature of the outside wall; the second step is the calculation of the convective heat flux at engine conditions, using the experimental heat flux and the Green functions engine-scalable technique. Inverse methodologies allow the determination of boundary conditions; in this case, the inner casing surface heat flux, based on measurements from outside of the system, which prevents aerothermal distortion caused by routing the instrumentation into the test article. The heat flux, retrieved from the inverse heat transfer methodology, is related to the rotor tip gap. Therefore, for a given geometry and tip gap, the pressure and temperature can also be retrieved. In this work, the digital filter method is applied in order to take advantage of the response of the temperature to heat flux pulses. The discrete Green's function approach employs a matrix to relate an arbitrary temperature distribution to a series of pulses of heat flux. In this procedure, the terms of the Green's function matrix are evaluated with the output of the inverse heat transfer method. Given that key dimensionless numbers are conserved, the Green's functions matrix can be extrapolated to engine-like conditions. A validation of the methodology is performed by imposing different arbitrary heat flux distributions, to finally demonstrate the scalability of the Green's function method to engine conditions. A detailed uncertainty analysis of the two-step routine is included based on the value of the pulse of heat flux, the temperature measurement uncertainty, the thermal properties of the material, and the physical properties of the rotor casing.

Uncertainty Quantification of Leakages in a Multistage Simulation and Comparison With Experiments

This paper presents a numerical study of the impact of tip gap uncertainties in a multistage turbine. It is well known that the rotor gap can change the gas turbine efficiency, but the impact of the random variation of the clearance height has not been investigated before. In this paper, the radial seals clearance of a datum shroud geometry, representative of steam turbine industrial practice, was systematically varied and numerically tested by means of unsteady computational fluid dynamics (CFD). By using a nonintrusive uncertainty quantification (UQ) simulation based on a sparse arbitrary moment-based approach, it is possible to predict the radial distribution of uncertainty in stagnation pressure and yaw angle at the exit of the turbine blades. This work shows that the impact of gap uncertainties propagates radially from the tip toward the hub of the turbine, and the complete span is affected by a variation of the rotor tip gap. This amplification of the uncertainty is mainly due to the low-aspect ratio of the turbine, and a similar behavior is expected in high pressure (HP) turbines.

Effects of Double-Leakage Tip Clearance Flow on the Performance of a Compressor Stage With a Large Rotor Tip Gap

Effects of a large rotor tip gap on the performance of a one and a half stage axial compressor are investigated in detail with a numerical simulation based on large eddy simulation (LES) and available particle image velocimetry (PIV) data. This paper studies the main flow physics, including why and how the loss generation is increased with the large rotor tip gap. The present study reveals that when the tip gap becomes large, tip clearance fluid goes over the tip clearance core vortex and enters into the next blade's tip gap, which is called double-leakage tip clearance flow. As the tip clearance flow enters into the adjacent blade's tip gap, a vortex rope with a lower pressure core is generated. This vortex rope breaks up the tip clearance core vortex of the adjacent blade, resulting in a large additional mixing. This double-leakage tip clearance flow occurs at all the operating conditions, from design flow to near stall condition, with the large tip gap for the current compressor stage. The double-leakage tip clearance flow, its interaction with the tip clearance core vortex of the adjacent blade, and the resulting large mixing loss are the main flow mechanism of the large rotor tip gap in the compressor. When the tip clearance is smaller, flow near the end wall follows more closely with the main passage flow and this double-leakage tip clearance flow does not happen near the design flow condition for the current compressor stage. When the compressor with a large tip gap operates at near stall operation, a strong vortex rope is generated near the leading edge due to the double-leakage flow. Part of this vortex separates from the path of the tip clearance core vortex and travels from the suction side of the blade toward the pressure side of the blade. This vortex is generated periodically at near stall operation with a large tip gap. As the vortex travels from the suction side to the pressure side of the blade, a large fluctuation of local pressure forces blade vibration. Nonsynchronous blade vibration occurs due to this vortex as the frequency of this vortex generation is not the same as the rotor. The present investigation confirms that this vortex is a part of separated tip clearance vortex, which is caused by the double-leakage tip clearance flow.

Mechanism study of performance enhancement in a subsonic axial flow compressor with recirculating casing treatment

The improvements in both compressor efficiency and stability have been observed in a subsonic axial flow compressor with recirculating casing treatment. The study aims to understand the underlying flow mechanisms of the beneficial effect on the compressor efficiency of recirculating casing treatment. The recirculating casing treatment was investigated experimentally and numerically. In the experiment, static pressure fluctuations over the rotor tip were measured with fast-response pressure transducers. Whole-passage time-accurate simulations were also implemented to help in understanding the flow details. The unsteadiness of double-leakage flow and its effect on the loss generated at the rotor tip were discussed in detail for the solid casing. The effect of recirculating casing treatment on the double-leakage flow related loss was subsequently investigated. The results indicate that the double leakage flow is a main loss source in the rotor tip region. The double-leakage flow is completely unsteady, which induces a cyclical fluctuation of the amount of loss generated in the rotor tip region for the solid casing. With the recirculating casing treatment installed, the high-energy jet created in recirculating loops changes the unsteady characteristics of double-leakage flow by getting it diffused over the rotor tip gap along the chordwise direction, which results in the loss fluctuation that is dominated by the high-energy jet rather than the unsteadiness of double-leakage flow. The double-leakage flow is also pushed downstream by the jet and the amount is reduced when it passes through the recirculating loops. The effect of recirculating casing treatment on the double-leakage flow is primarily responsible for the improvement in the compressor efficiency.

Experimental study of self-recirculating casing treatment in a subsonic axial flow compressor

Parametric studies of recirculating casing treatment were experimentally performed in a subsonic axial flow compressor. The recirculating casing treatment was parameterized with injector throat height, injection position, and circumferential coverage percentage. Eighteen recirculating casing treatments were tested to study the effects on compressor stability and on the compressor overall performance at three blade speeds. The profiles of recirculating casing treatment were optimized to minimize the losses generated by air recirculation. In the experiment, the stalling mass flow rate, total pressure ratio, and adiabatic efficiency of the compressor were measured to study the steady-state effects on the compressor performance of recirculating casing treatments, and static pressure disturbances on the casing wall were monitored to study the influence on stall dynamics. Results indicate that both the compressor stability and overall performance can be improved through recirculating casing treatment with appropriate geometrical parameters for all the test speeds. The influence on stall margin of one geometric parameter often depends on the choice of others, i.e. the interaction effects exist. In general, the recirculating casing treatment with a moderate injector throat and a large circumferential coverage is the optimal choice to enhance compressor stability. The injector of recirculating casing treatment should be placed upstream of the blade tip leading edge and the injector throat height should be lower than four times the rotor tip gap for the benefits of compressor efficiency. At 71% speed, the blade tip loading is decreased through recirculating casing treatment at the operating condition of near peak efficiency and increased near stall. Moreover, the outlet absolute flow angle is reduced in the tip region and enhanced at lower blade spans for both operating conditions. The stall inceptions are not changed with the implementation of recirculating casing treatment for all the test speeds, but the stall patterns are altered at 33% and 53% speeds, i.e. the stall with two cells is detected in the recirculating casing treatment compared with the solid casing with only one stall cell.

Mechanism of Nonsynchronous Blade Vibration in a Transonic Compressor Rig

This paper presents a numerical study on blade vibration for the transonic compressor rig at the Technische Universität Darmstadt (TUD), Darmstadt, Germany. The vibration was experimentally observed for the second eigenmode of the rotor blades at nonsynchronous frequencies and is simulated for two rotational speeds using a time-linearized approach. The numerical simulation results are in close agreement with the experiment in both cases. The vibration phenomenon shows similarities to flutter. Numerical simulations and comparison with the experimental observations showed that vibrations occur near the compressor stability limit due to interaction of the blade movement with a pressure fluctuation pattern originating from the tip clearance flow. The tip clearance flow pattern travels in the backward direction, seen from the rotating frame of reference, and causes a forward traveling structural vibration pattern with the same phase difference between blades. When decreasing the rotor tip gap size, the mechanism causing the vibration is alleviated.

System identification modeling and unstable behavior of the dynamics of flows within the tip region of an axial compressor blade passage

In recent years, the correlation coefficient of pressure data from the same blade passage in an axial compressor unit has been used to characterize the state of flow in the blade passage. In addition, the correlation coefficient has been successfully used as an indicator for active control action using air injection. In this work, the correlation coefficient approach is extended to incorporate system identification algorithms in order to extract a mathematical model of the dynamics of the flows within a blade passage. The dynamics analyzed in this research focus on the flow streams and pressure along the rotor blades as well as on the unsteady tip leakage flow from the rotor tip gaps. The system identification results are used to construct a root locus plot for different flow coefficients, starting far away from stall to near stall conditions. As the compressor moves closer to stall, the poles of the identified models move towards the imaginary axis of the complex plane, indicating an impending instability. System frequency data is captured using the proposed correlation based system identification approach. Additionally, an oscillatory tip leakage flow is observed at a flow coefficient away from stall and how this oscillation changes as the compressor approaches stall is an interesting result of this research. Comparative research is analyzed to determine why the oscillatory flow behavior occurs at a specific sensor location within the tip region of the rotor blade.

Experimental investigation of the effect of rotor tip gaps on 3D separating flows inside the stator of a highly loaded compressor stage

Flow separations in compressor blade passages are common and can cause significant flow blockage and loss production. This paper investigates experimentally the three-dimensional (3D) flow separations in a highly loaded low-speed large-scale compressor facility. Oil flow visualizations, stereoscopic particle image velocimetry (SPIV), and five-hole probe measurements are conducted at certain conditions from the near-chock to near-stall condition along the compressor operating line. The 3D separation and vortex flow structures in the stator at different operating conditions are analyzed. By changing the size of the rotor tip gap, six groups of oil-flow pictures are obtained to study the effect of the rotor gaps on the 3D separating flows in the downstream stator. The variation of the corner separation scale is almost linear with the rotor tip gap size. Along the compressor operating line, four typical 3D flow structures are found inside the stator passage. Between the second typical 3D flow structure and the third typical 3D flow structure, an unstable stage exists on the compressor operating line; during this stage, the hub corner separation becomes an open separation from the closed type. A smaller rotor tip gap corresponds to an earlier unstable stage. Finally, a critical rotor tip blockage state usually existed for the transform of corner separation types at a certain rotor tip gap configuration. This discovery is valuable for the study of multistage compressor matching problems at off-design conditions.

Analysis of the Heat Transfer Driving Parameters in Tight Rotor Blade Tip Clearances

Turbine rotor tips and casings are vulnerable to mechanical failures due to the extreme thermal loads they undergo during engine service. In addition to the heat flux variations during the engine transient operation, periodic unsteadiness occurs at every rotor passage, with amplitude dependent on the tip gap. The development of appropriate predictive tools and cooling schemes requires the precise understanding of the heat transfer mechanisms. The present paper analyses the nature of the overtip flow in transonic turbine rotors running at tight clearances and explores a methodology to determine the relevant flow parameters that model the heat transfer. Steady-state three-dimensional Reynolds-averaged Navier–Stokes (RANS) calculations were performed to simulate engine-like conditions considering two rotor tip gaps, 0.1% and 1%, of the blade span. At tight tip clearance, the adiabatic wall temperature is no longer independent of the solid thermal boundary conditions. The adiabatic wall temperature predicted with the linear Newton's cooling law was observed to rise to unphysical levels in certain regions within the rotor tip gap, resulting in unreliable convective heat transfer coefficients (HTCs). This paper investigates different approaches to estimate the relevant flow parameters that drive the heat transfer. A novel four-coefficient nonlinear cooling law is proposed to model the effects of temperature-dependent gas properties and of the heat transfer history. The four-parameter correlation provided reliable estimates of the convective heat transfer descriptors for the 1% tip clearance case, but failed to model the tip heat transfer of the 0.1% tip gap rotor. The present study allows experimentalists to retrieve information on the gap flow temperature and convective HTC based on the use of wall heat flux measurements. The use of nonlinear cooling laws is sought to improve the evaluation of the rotor heat transfer data while enhancing the understanding of tight-clearance overtip flows.