Gas turbine designers demand accurate predictions of metal temperature to ensure acceptable operating life of components experiencing high thermal stress. Rotor-stator cavities ingest hot mainstream gas through rim seals when inadequately purged with relatively cool air bled off the compressor. Superfluous use of purge, and any associated windage increase, creates a parasitic loss in overall efficiency. Shear interaction caused by the difference in swirl between the purge and mainstream flow is a principal driver for ingestion; pre-swirled purge flow has the potential to alter the swirl gradient. This paper presents the first assessment of purge conditioning in a downstream cavity. An experimental campaign was conducted in a new aeroengine representative 1.5-stage test facility designed to facilitate expedient changeover of modular components in the downstream stator assembly. Purge flow in the downstream cavity was supplied through a series of angled injectors contained in a single component at mid-radius. Three co-swirled injection angles were tested. Measurements of CO2 gas concentration, static pressure and swirl were taken in the cavity to examine the relationship between purge-mainstream swirl gradient and ingress downstream of a rotor blade. The aero-engine designer must balance caution when employing pre-swirl to reduce disc windage; co-swirled purge increased the purge-mainstream swirl gradient and subsequently increased shear-driven ingestion.

Next-generation aero-engine compressors will operate with overall pressure ratios exceeding 70:1. This will require shorter compressor blades, presenting a challenge to the designer when predicting tip clearance and efficiency. Buoyancy-induced flow within co-rotating compressor discs drives the heat transfer that determines rotor expansion and the resulting blade-tip clearance. This inherently unstable flow is influenced by the radial temperature distribution of the discs, rotational speed, as well as enthalpy and momentum exchange with an axial throughflow of cooled air at low radius. Due to the rotation of the engine compressor, this throughflow may become swirled, altering the temperature, mass exchange, and swirl within the rotating cavity. The University of Bath Compressor Cavity Rig has been adapted to introduce preswirl into the axial throughflow by passing it through rotating holes. The effects of inlet swirl have been characterized in terms of Rossby and Reynolds numbers. Measurements of disc temperature, shroud heat flux, and unsteady pressure in the rotating frame of reference are used to quantify the effects of ingestion (entrainment) of fluid into the cavity. The unsteady dynamics and rotation of the core relative to the disc have been measured in both the stationary and rotating frames of reference with consistent results. A single correlation between shroud Nusselt and Grashof numbers has been established, effectively capturing the impact of swirl, Rossby number, and free convection.

The buoyancy-driven flow inside compressor cavities is three-dimensional, unsteady and features a large range of time- and length-scales. The temperature and rotation of the fluid core of the cavity are influenced by an exchange (and recirculation) of enthalpy and momentum with an axial throughflow of cooling air at low radius. The complexity of the flow and conjugate nature of the heat transfer to the discs creates a challenge for the aero-engine designer when calculating thermal stresses, radial expansion, and blade-tip clearances. This paper presents a low-order model to predict the radial variation of disc and fluid-core temperatures, and the mass exchange (entrainment) to the rotating cavity. Fundamental physical principles and experimental data are used to create a single set of Rayleigh-Grashof correlations for heat transfer and radial mass flow of buoyant plumes. The model is applied to eleven test cases from experimental rigs at Bath, Dresden and Sussex, each with unique instrumentation, thermal boundary conditions, geometries and throughflow swirl. Empirical correlations for exchange and recirculation mass flow were determined for each rig using a common theoretical methodology. The model captures the heat and mass transfer characteristics with accuracy quantified relative to experimental data. New experimental data from the Bath Compressor Cavity Rig is used to validate the model under conditions of asymmetrical heating, demonstrating the effects associated with the axial gradient of temperature in the compressor are captured appropriately. The consistent agreement with experimental data and correlation methodology demonstrates a robust framework appropriate for application to thermo-mechanical design codes in the aero-engine.

The Secondary Air System in an aero-engine uses relatively cool purge from the compressor to limit the ingress of hot annulus gases into vulnerable turbine cavities through rim seals. Superfluous use of purge is inefficient, while insufficient use leads to thermal degradation of highly-stressed turbine components. This study establishes a predictive design tool to fully characterise a rim seal across the performance envelope of the engine. Physically-informed low-order models are important in the engine design process. The Ingress Wave Model uses a single, empirically-correlated parameter to physically link the shear-driven unsteadiness in the cavity with the swirl in the annulus. In this paper, newly-collected and existing experimental data from three facilities and rim seal geometries demonstrated that this unsteadiness is a linear function of the annulus swirl, with the superposition of purge creating a weak, secondary effect. The novel introduction of the linear correlation enables predictions of sealing effectiveness across the entire engine operating range from just two data points. A larger set of data is shown to improve the accuracy and robustness. Here the method is validated by data collected at low technology readiness level (TRL). The methodology could be applied to data collected from a high TRL demonstrator engine or computational fluid dynamics. This will reduce the number of demonstrator experiments (and associated costs) during design iterations. This paper provides an original scaling methodology in the practical context of the engine design process, including the effects of density. Aero-engines operate with a significant purge-mainstream density ratio (DR), due to differences in the temperatures of the two streams. The methodology incorporates predictions of DR and is further validated against data collected at DR = 1 and 1.5. The model demonstrates that neglecting DR will provide significantly underpredicted rim seal performance.

The efficiency of aero-engines is linked to increased turbine entry temperature and a secondary air system that protects vulnerable components under high thermal stresses and metal temperatures. Purge (or sealing) air from the compressor is used to limit the ingress of hot mainstream annulus gases into rotor-stator cavities in the high-pressure turbine. Accurately predicting ingress, and understanding conditions under which it is amplified, is a significant challenge for the engine designer. Experimental data gathered from a 1.5-stage turbine facility and a mathematical, physics-informed model are used to link the rotation of large-scale structures (instabilities) near the rim seal with amplified ingress. The Ingress Wave Model identifies the swirl of cyclonic-anticyclonic vortex pairs (instabilities) in the cavity as the transport mechanism for ingress. The intensity of these unsteady rotating structures is maximised if the circumferential pressure field in the cavity is synchronised (hence superposition) to that in the annulus. Cross-correlation of unsteady pressure measurements in the cavity forward of the rotor revealed this synchronisation was to the pressure field caused by downstream rotating blades. In the aft cavity, this synchronisation was in the stationary frame of reference and associated with the downstream vanes. The effects of amplified ingress are shown to be significant and exist in turbine rigs featuring a wide range of blade and vane counts. In terms of new knowledge and originality, the synchronisation to the pressure field provides the first explanation of this important physical mechanism. A criterion for the engine designer to avoid this phenomenon is proposed.

In an aero-engine compressor, co-rotating discs form cavities that interact with an axial throughflow of secondary air at low radius. In the high-pressure (HP) compressor the shroud is hotter than the throughflow (directed downstream to the turbine) and the radial temperature gradient creates buoyancy-induced flow at Grashof numbers similar to 10(13). Such flows can be unstable and typically take the form of counter-rotating vortex pairs separated by radial hot and cold plumes. However, in low pressure (LP) and intermediate pressure (IP) compressors the secondary air is directed upstream. In this inverse scenario, the axial throughflow is hotter than the compressor discs, reversing the disc temperature gradient and eliminating the fundamental driver for buoyancy. Despite its practical application and importance, this inverse scenario has not been previously investigated. The University of Bath Compressor Cavity Rig has been uniquely designed to simulate such flows, measuring temperature and unsteady pressure in the frame of reference of the rotating discs. Bayesian and spectral analysis have determined the radial distribution of disc heat flux, as well as the asymmetry of the rotating vortex structures and their slip relative to the discs. Unexpectedly, the new data reveal the flow structure in cavities with positive and inverted temperature differences are fundamentally similar (albeit with reversed radial-temperature profiles). Isothermal cases identified a critical Rossby number (Ro), above which the flow structure in the cavity was dominated by a toroidal vortex. At sub-critical Ro, the flow structure for the inverted temperature gradient continued to be governed by buoyancy due to disc heat transfer. Momentum exchange with the axial throughflow and the gradient of circumferential pressure combine to vary the slip and vortex symmetry. This paper provides the first data and analysis of flow and heat transfer during inverse throughflow conditions in LP and IP compressors. The new insights are of importance for the determination of the thermal stresses in discs, engine life, compressor blade clearance and efficiency.

The next generation of aeroengines will feature compressors with increasing pressure ratios and smaller engine cores. Maintaining high efficiencies will require increased sensitivity to reduced blade tip clearances, governed by strong buoyancy-induced flow and heat transfer within the rotating cavities formed by the discs to which the blades are attached. The inherently unsteady flow within these cavities is three-dimensional and unstable. Thermal stresses in the discs are governed by forced and natural convection across large differences in temperature, conjugate heat transfer, centrifugal forces, and disrupted by mass exchange between the core air and an axial cooling throughflow at low radius. The thermo-fluid-dynamics has further complexity during accelerations or decelerations in aeroengine transients. The engine design process requires expedient and reliable aerothermal models to predict the transitory temperatures of the discs, and hence the thermal growth of the rotor and tip clearance. This paper presents, for the first time, a theoretical model to predict compressor cavity transient heat transfer and temperatures from first principles. The reduced-order model was created in close partnership with an experimental programme using an innovative rig designed specifically to explore buoyancy-induced flow in compressor cavities. Unsteady pressure, temperature and heat flux data were collected in the rotating frame of reference under controlled boundary conditions for two engine-representative open-cavity configurations. The predicted temperature, mass flow and heat flux results were consistent with measured values within experimental uncertainty as shown by RMSE analysis. The research has identified sub- and super-critical flow regimes governed by the Rossby number and enthalpy exchange with the axial-throughflow. Fundamental insight has been established, including the presence of large-scale structures formed from buoyancy-induced convection as well as that induced purely by throughflow interaction. In collaboration with Rolls-Royce, this study has been framed in the practical context of providing expedient solutions appropriate for transient thermo-mechanical codes during the engine design process.

Unraveling the flow physics pertaining to hot gas ingress in turbines is crucial in enabling designers to realize global decarbonization targets in aerospace. A turbine rim seal is fitted at the periphery of the rotor-stator cavity to minimize the ingress of annulus gas, which detrimentally affects cycle's efficiency. The inherent unsteadiness in rim seal flows, arising from shear gradients between contiguous flow paths, introduces a consequential, yet presently unestablished, influence on sealing characteristics. A single-stage axial turbine facility in conjunction with an aeroengine architecture is employed to assess the steady and unsteady sealing characteristics of a range of industrially relevant rim seals. Time-averaged measurements of gas concentration and swirl, acquired over a range of flow coefficients (C-F), exhibited an inverse relationship between sealing performance and the purge-mainstream swirl difference (Delta beta). Spectral analysis of unsteady pressure signals revealed an associated unsteadiness, induced by the strength of the annulus-cavity interaction. Across all C-F, a low-frequency harmonic range consistently displayed proportionality between spectral activity and Delta beta. Thus, a relationship between steady and unsteady characteristics was established. Examining a series of rim seal configurations with varying radial clearances signified that sealing performance was predominantly influenced by the radially outermost clearance. The configurations exhibiting superior performance presented heightened spectral activity, ascribed to an increased radial purge mass flux and establishing a definite relationship with concurrent steady measurements.

As new aeroengine architectures move to larger diameter fans and rotors, the associated increase in weight will need to be counterbalanced by lighter, more compact, and more efficient low-pressure turbines (LPT). Efficiency gains in LPTs can be achieved by reducing the losses associated with the shroud leakage flows. Flow control studies on the topic have traditionally focused on reducing the mixing loss, which constitutes a considerable proportion of the total loss. Nonetheless, increasing engine speeds are driving additional gains obtained by also targeting the reduction of windage losses. Developing a flow control solution with the dual objective of reducing over-tip cavity mixing and windage losses has not previously been attempted. This is a challenge due to conflicting flow control requirements and geometric constraints. Reducing windage loss generally requires increasing the swirl-ratio of the leakage flow, while reducing mixing loss requires reducing this ratio to match that of the main gas path. The current work proposes a novel flow control solution to successfully achieve this purpose through the emerging technology of additive manufacturing. The successful flow control concept was developed through numerical simulation, printed using an additive manufacturing process, and validated in a purpose-built rig. Experiments and computations were consistent with a cumulative reduction in cavity windage of 16%. The FCC is estimated to increase the mechanical efficiency of the turbine stage in isolation by 1%.

Significant density ratios arise in a gas turbine due to severe temperature gradients between the hot mainstream gases leaving the combustor and the superposed purge flow injected from the secondary air system. Engineers seek to minimise the ingestion of hot annulus gas through the rim seal at the periphery of the turbine wheel-space to maximise component life while continuing to increase the turbine entry temperature in pursuit of optimised thermodynamic cycle efficiency. The majority of experimental ingestion facilities assess sealing performance at a near-unity purge–mainstream density ratio which negates the impact of this significant contributor to ingestion. This study investigates the impact of the density ratio on the fluid mechanics across the rim seal of a single-stage turbine facility. The results demonstrate that the purge–mainstream density ratio is a crucial consideration when designing the rim seal architecture, particularly with the transition to alternative fuels which have the potential to augment the temperature gradient. A density-affected region at the intermediate superposed purge flows is identified where the non-unity density ratio has the greatest impact on outer cavity swirl and sealing effectiveness. Furthermore, unsteady pressure spectra in this region exhibit a suppression of the low-frequency spectral band as the density ratio is increased, highlighting a causal link between unsteadiness and ingress.

Next generation aeroengines will operate at ever-increasing pressure ratios with smaller cores, where the control of blade-tip clearances across the flight cycle is an emerging design challenge. Such clearances are affected by the thermal expansion of the compressor discs that hold the blades, where acute thermal stresses govern operating life. The cavities formed by co-rotating discs feature a heated shroud at high radius and cooler cobs at low radius. A three-dimensional, unsteady and unstable flow structure is induced by destabilising buoyancy forces. The radial distribution of disc temperature is driven by a conjugate heat transfer at Grashof numbers of order 1013. Such flows are further influenced by the heat and mass exchange with an axial through-flow of cooling air at low radius, where the interaction depends on the Rossby number and separation of the disc cobs. This paper is the first to study the effect of cob separation ratio on mass and heat exchange for compressor cavities. A model is developed to predict the cavity-throughflow interaction, and disc and fluid-core temperatures. The judicious use of a physics-based methodology provides reliable, reduced-order solutions to the complex conjugate problem, thereby making it appropriate for practical engine thermo-mechanical design. The model is validated by detailed experimental measurements using the Bath Compressor Cavity Rig, where variable disc cob spacings were investigated over a range of engine-representative conditions. The unsteady pressure measurements collected in the frame of reference of the rotating discs reveal new insight into the fundamentally aperiodic nature of the flow structure. This new understanding of heat transfer informs an expedient reduced-order model and enables more efficient design of future high pressure-ratio aeroengines.

Aeroengines operate with a cooling flow (purge) at a significant purge-mainstream density ratio (DR), which is principally created by the differences in temperatures of these two streams. This paper will show there is a profound influence of DR on ingress, purge flow rates, and sealing effectiveness - all crucial to the superordinate aim of achieving a high thermodynamic efficiency for the engine. A new theoretical (low-order) model is introduced to enable the engine designer to flexibly predict the required purge to prevent ingress over a range of typical operating conditions. The Ingress Wave Model is based on the physical principle that unsteadiness, in the form of large-scale rotating instabilities, forms a circumferential pressure gradient driving fluidic motion against the Coriolis force. The shear created by the difference in tangential momentum between adjacent flow streams is assumed to be the primary mechanism in the process. This allows a set of equations to be derived from dimensional analysis and the assumption that flow entrainment is a function of the relative egress momentum and ingress density. The model is validated against data collected at both DR = 1 and 1.52, with good quantitative agreement across a range of purge and annulus flow conditions. Typical engine design practice exploits information captured in experimental rigs operating in benign conditions at low technology readiness level (TRL) and DR = 1. The new model is used to scale such data collected from six experimental facilities to the density ratios expected in current state-of-the-art (DR = 1.5) and future (DR = 2) engines. The result is a requirement for significantly reduced purge, with profound practical implications for the engine designer, in particular for future engines which operate at higher purge-mainstream density ratios.

The windage torque on rotational walls has negative effect on the performance of the low pressure turbine. In this paper, three novel flow control concepts (FCCs) were proposed to reduce the windage torque within a turbine stator well, with upstream and downstream cavities connected by an interstage labyrinth seal. The swirl and flow pattern inside a reference turbine cavity was first investigated and the potential locations for the FCCs were identified using numerical simulations. FCC1 was a circumferential row of leaned deflectors downstream of the labyrinth seal. FCC2 was a set of deflector vanes and platform to optimize the ingress swirl at high radius in the upstream cavity. FCC3 combined the two flow concepts and the superposition resulted in a stator well windage torque reduction of 70% when compared to the baseline design. The FCCs also showed performance benefits at off-design conditions and over a range of secondary flow rates to the cavity. In Part 2 [1], the numerical analysis and performance of the FCCs are validated in an experimental rig, using additively-manufactured components.

Recent experimental and computational research has demonstrated unsteady large-scale flow structures rotating in the rim seal clearance of aero-engine turbines. This intrinsic unsteadiness dominates the physical mechanism for the detrimental ingress of hot gases. Existing theoretical models of ingress lack this physical connection; nor can they predict the change in pressure across rim seals, important to the engine designer setting a superposed sealing flow to suppress ingress. This paper presents a new low-order model, named the Ingress Wave Model, relating ingress to the strength of the large-scale flow structures. Sinusoidal functions are used to represent the radial and circumferential velocities within the structures. For a given superposed sealing flow rate, a sealing effectiveness is determined by the non-dimensional amplitude of the sinusoidal function for the radial velocity, ΦA, an empirical term that depends on the seal geometry as well as the annulus and wheel-space flow conditions. The circumferential and radial velocity are linked via the equations for continuity and angular momentum, yielding the radial pressure drop across the seal. Using appropriate empirical values of ΦA, the model is validated over a wide range of experimental data from the literature. It is also demonstrated that the model is a practical tool in the preliminary design process, estimating the degree of ingress under different operating conditions.

Ingress is the penetration of hot mainstream gas into the rotor-stator wheel-space formed between adjacent discs; a rim seal is installed at the periphery of the wheel-space. Purge flow is bled from the compressor and re-introduced in the turbine to reduce, or in the limit prevent, ingress. This study presents a unique, concomitant experimental and turbulence-resolved numerical investigation of ingress in an aeroengine rim seal geometry, with leakage flow. Experimental modelling is conducted in the University of Bath’s 1-stage turbine test facility. Measurements of gas concentration, pressure and swirl were used to assess the performance of the rim seal. A parallel study using Improved Delayed Detached Eddy Simulations (IDDES) was used to generate time-averaged and time-resolved flow-fields, enabling direct comparison with experimental data. The aeroengine architecture conformed to classical rim seal mechanics whereby the effectiveness level increased with purge flow. The inner wheel-space exhibited Batchelor-type flow; in the outer wheel-space, the effectiveness was nonaxisymmetric and synchronised in accordance with the local radial velocity field. Utilisation of a DES TKE multiplier demonstrated regions where increased turbulence resolution was required to resolve the appropriate scale of turbulent eddies. IDDES computations were found to accurately capture the radial distributions of pressure, swirl and effectiveness, both in the absence and superposition of leakage flows. The IDDES approach exhibits significantly superior agreement with experiments when compared to previous studies employing an Unsteady Reynolds Averaged Navier-Stokes (URANS) methodology.

Minimizing the losses within a low-pressure turbine (LPT) system is critical for the design of next-generation ultra-high bypass ratio aero-engines. The stator-well cavity windage torque can be a significant source of loss within the system, influenced by the ingestion of mainstream annulus air with a tangential velocity opposite to that of the rotor. This paper presents experimental and numerical results of three carefully designed Flow Control Concepts (FCCs)—additional geometric features on the stator surfaces, which were optimized to minimize the windage torque within a scaled, engine-representative stator-well cavity. FCC1 and FCC2 featured rows of guide vanes at the inlet to the downstream and upstream wheel-spaces, respectively. FCC3 combined FCC1 and FCC2. Superposed flows were introduced to the upstream section of the cavity, which modelled the low-radius coolant and higher radius leakage between the rotor blades. In addition to torque measurements, total and static pressures were collected, from which the cavity swirl ratio was derived. Additional swirl measurements were collected using a five-hole aerodynamic probe, which traversed radially at the entrance and exit of the cavity. A cavity windage torque reduction of 55% on the baseline (which has no flow control) was measured for FCC3, at the design condition with superposed flow. For this concept, an increase in the cavity swirl in both the upstream and downstream wheel-spaces was demonstrated experimentally and numerically. With increasing superposed flow, the contribution of FCC1 surpassed FCC2, due to more mass flow entering the downstream wheel-space across the rotor fins (passing FCC1), and less ingestion from the annulus into the upstream wheel-space (passing FCC2). The torque changes from the concepts are explained using the fluid dynamic evidence from experimental swirl measurements and computational simulations. The simulations allow translation to engine-operating conditions and practical information to the engine designer.

Robust methods to predict heat transfer are vital to accurately control the blade-tip clearance in compressors and the radial growth of the discs to which these blades are attached. Fundamentally, the flow in the cavity between the co-rotating discs is a conjugate problem: the temperature gradient across this cavity drives large-scale buoyant structures in the core that rotate asynchronously to the discs, which in turn governs the heat transfer and temperature distributions in the discs. The practical engine designer requires expedient computational methods and low-order modelling. A conjugate heat transfer methodology that can be used as a predictive tool is introduced here. Most simulations for rotating cavities only consider the fluid domain in isolation and typically require known disc temperature distributions as the boundary condition for the solution. This paper presents a novel coupling strategy for the conjugate problem, where unsteady Reynolds Averaged Navier-Stokes (URANS) simulations for the fluid are combined with a series of steady simulations for the solid domain in an iterative approach. This strategy overcomes the limitations due to the difference in thermal inertia between fluid and solid; the method retains the unsteady flow features but allows a prediction of the disc temperature distributions, rather than using them as a boundary condition. This approach has been validated on the fundamental flow configuration of a closed co-rotating cavity. Metal temperatures and heat transfer correlations predicted by the simulation are compared to those measured experimentally for a range of engine-relevant conditions.

Accurate prediction of heat transfer in compressor cavities is crucial to the design of efficient and reliable aircraft engines. The heat transfer affects the thermal expansion of the compressor rotor and, in turn, the tip clearance of the compressor blades. This paper presents a novel, physically-based predictive theoretical model of heat transfer and flow structure in an open compressor cavity, which can be used to accurately calculate disc temperatures. The radially higher region of the cavity is dominated by buoyancy effects created by the temperature difference between the hot mainstream flow and the axial throughflow used to cool the turbine. Strong interaction between the air in the cavity and this throughflow creates a mixing region at low radius. For a given geometry, the heat transfer and flow physics are governed by four parameters: the rotational Reynolds number Reϕ, the buoyancy parameter βΔT, the compressibility parameter χ, and the Rossby number Ro. The model quantifies both the buoyancy- and throughflow-induced mass and heat transfer, producing a reliable prediction of the disc and air temperatures. The model takes into account a two-fold effect of the throughflow: being entrained into the cold radial plumes directly and creating a toroidal vortex in the radially lower region of the cavity. The exchange of mass between the cavity and throughflow is related to the mass flow rate in the radial plumes in the buoyancy-induced region, considering the effect of flow reversal at low Ro. The model is validated using data collected in the Bath Compressor Cavity Rig and can be incorporated in engine design codes to robustly compute the thermal stress and expansion of the compressor rotor, contributing to more efficient engine designs.

As part of the drive towards achieving net-zero flight by 2050, the development of ultra-high bypass ratio engines will place greater importance on the efficiency of the low-pressure turbine. Stator-well cavities can be a significant source of loss within this system, primarily due to the ingestion of annulus air, which can rotate at a speed less than that of the disc. This paper presents the results of a joint experimental/numerical campaign to relate windage torque to the flow physics in a scaled, engine-realistic stator-well geometry with superposed flows. The primary variables were the pressure ratio across the stator row and the swirl ratio at inlet to the cavity. This is the first paper to relate simultaneous torque and swirl measurements in a stator-well cavity. The numerical simulations have also provided insight into the fluid dynamic behaviour. A reduction in cavity windage torque with superposed flow rate was shown, consistent with the increase in swirl measured by experiments and predicted by computations. The pre-swirled superposed flows reduce the ingestion of negatively swirling fluid from the annulus, effectively increasing the swirl ratio of the flow in the cavity towards that of the rotor. This reduces windage losses; at the design condition with superposed flow, the cavity windage torque reduced by 60% of that of the reference case. Sealing effectiveness increased with superposed flow rate. This was demonstrated through gas concentration measurement/simulations, a reduced radial flow velocity into the cavity and a reduced mass flow rate across the interstage seal. The application of a turbulent transport model showed that ingestion could be explained by shear-driven diffusion. Ingestion increased as inlet swirl ratio reduced, while the pressure ratio was found to have a negligible influence on windage moment coefficient, swirl ratio and sealing effectiveness. In addition to providing fluid dynamic insight, the gas concentration results can be used for validating thermomechanical models. The results in this paper will be of practical interest to the engine designer who wishes to scale information to engine-operating conditions.

During transient gas turbine operating cycles, the radial clearance between the rotating blades and the stationary casing is governed by the thermal stresses and radial growth of the compressor discs. Robust and accurate prediction of the temperature and heat transfer inside compressor cavities is critical to controlling these clearances and the design of reliable and efficient engines. This paper presents a fully predictive model to determine disc temperatures and shroud heat fluxes for a closed disc cavity (rotating cylindrical annulus), simulating an engine compressor. This is the first published reduced-order modelling of compressor cavities during engine transients in literature and the most detailed publication of transient experimental data in this context. Due to the conjugate nature of the heat transfer, a conduction model of the compressor disc and a model of the flow and heat transfer in the cavity between the co-rotating discs are included. The conduction was computed using a two-dimensional finite element solver, and the disc heat transfer was calculated assuming conductive laminar Ekman layers on the disc surfaces. Correlations for free convection on flat plates are used to model the heat transfer on the inner and outer radii of the rotating cavity. The transient core temperatures were calculated in a quasi-steady manner, with the heat transfer at every time step predicted assuming zero net heat flow. The model was validated using new transient data collected from the Compressor Cavity Rig at the University of Bath. The model shows consistent agreement with all experimental cases, providing insight into the conjugate relationship between the disc and shroud heat transfer, revealing that disc heat transfer is sustained long after there is no convective heat transfer from the shroud. This approach has direct application to practical thermo-mechanical codes, contributing to the design of next-generation, net-zero aerospace architectures.