Efficient cooling of the trailing edge channel is a critical aspect of internal cooling design for turbine blades. In this study, a trapezoidal (or wedge-shaped) channel geometry with lateral overflow slots is developed, incorporating staggered inclined ribs and pin-fins to form a composite cooling structure. The investigated Reynolds number (Re) and rotation number (Ro) ranges are 8000-24000 and 0-0.83, respectively. By adjusting the density ratio (DR = 0.07-0.14), the buoyancy number (Buo) is maintained within the range of 0-2.01. Experiments are carried out using the copper plate method, with thin-film thermal resistors providing a constant heat flux. Numerical simulations are conducted based on the SST model. Results indicate that the rib-pin configuration enhances both overall and local heat transfer performance in the inner region, with staggered ribs particularly improving heat transfer on the leading wall by 19.8%. However, compared to pin-fins, ribs exhibit a notable decline in maintaining heat transfer efficiency under rotational conditions, with reductions exceeding 50%. Adjusting DR to generate varying Buo values remains a feasible approach for investigating buoyancy effects under the current test facility. Mathematical decoupling of Ro and Buo reveals that heat transfer within the inner region is predominantly governed by the Coriolis force, whereas the outer region is affected by both the Coriolis force and buoyancy. Spanwise average heat transfer correlations are proposed with a fitting accuracy of +/- 15%. Numerical simulation results demonstrate that modifying the radius ratio (rin/Dh) to alter Buo represents another effective method, facilitating a more comprehensive understanding of the influence of buoyancy on internal flow and heat transfer within turbine blades.

Physics-informed neural networks (PINNs) approximate solutions of partial differential equations by minimizing residual losses evaluated at collocation points. The collocation sampling rule therefore defines a finite residual estimator and can influence the optimization trajectory, not only the terminal error. Existing comparisons of residual-adaptive samplers often rely on final or oracle-selected best-checkpoint errors, which can conceal checkpoint sensitivity, seed variability, late regression, and sampling overhead. This paper develops a trajectory-resolved evaluation framework for residual-adaptive PINN sampling. The framework defines fixed final and last-window reports, exact-error oracle diagnostics, residual-validation-based non-oracle checkpoint interfaces where validation trajectories are available, instability and time-to-threshold summaries, paired seed-wise comparisons, protocol-induced rank sensitivity, sampling diagnostics, and wall-clock cost accounting. The framework is applied to random residual resampling, a residual-focused adaptive branch (RAD-strong), importance-corrected Weighted-RAD, tempered and scheduled correction variants, and a scoped residual-refinement reference across smooth, sharp-gradient, reaction-diffusion, high-dimensional, and sample-starved benchmark roles. The results show problem-dependent effects of importance correction: it is favorable for Burgers, provides positive 10-seed evidence for a sample-starved four-dimensional dimension-flexible heat (DF-heat) case with Nf = 500, acts as a stabilizing tradeoff in the Original 4D DF-heat case, and is less effective than residual focus in the Allen–Cahn setting. Schedule variants and residual-refinement references are most informative as trajectory diagnostics. Overall, the framework supports seed-aware, checkpoint-aware, validation-aware, and cost-aware evaluation of adaptive PINN sampling beyond final-error tables.

Efficient cooling of turbine blades is one of the key technologies for enhancing the performance and service life of aero engines. The turbine blade cooling structure consists of interconnected external film holes and internal cooling passages, which jointly cool and protect the blades. In previous studies on turbine blade internal cooling channels, the film hole outflow effect on flow dynamics and heat transfer has often been overlooked. This study employs numerical simulation to investigate how film hole outflow affects flow heat transfer within an internal cooling channel under both static and rotating conditions. The internal cooling channel is simplified as a square cross-section straight duct with outflow holes located only on the trailing and leading walls. The investigated Reynolds numbers are 15,000 and 30,000, the rotation number ranges from 0 to 0.44, the film hole outflow ratios vary between 0 and 1. The results indicate that the presence of film hole outflow disrupts the boundary layer on the downstream wall of the film holes, thus enhancing heat transfer significantly within the channel. The outflow from the film holes causes a gradual decrease in flow rate along the channel, resulting in low heat transfer regions downstream. Rotation weakens the impact of film hole outflow. The Coriolis force induced by rotation causes main stream to shift toward channel trailing wall, enhancing heat transfer while reducing spread difference of wall temperature caused by film hole outflow. For the leading wall, the deviation of main stream weakens the destructive effect of film hole outflow on boundary layer, thereby diminishing heat transfer. The results presented in this study reveal the mechanisms by which film hole outflow affects flow heat transfer within an internal cooling channel, providing valuable insights for optimizing turbine blade cooling structures.

Real-time optimization of aircraft thermal management systems (TMS) is constrained by the computational cost of conventional approaches. Solving the existing simulating model requires several seconds, and optimization through evolutionary algorithms based on above model further extends this by at least several minutes. None are viable for onboard real-time applications. Consequently, developing the rapid and accurate determination of TMS strategy has become a pivotal challenge. This study firstly proposes a real-time regulation strategy to reduce optimization time to the millisecond scale. Specifically, an onboard dynamic simulation model is developed that resolves only the heat transfer processes among the heat loads and heat sinks, and eliminates the pressure-balance iteration. The computation time is shortened from the second scale to real-time levels. Embedding above model, a multi-objective decision optimization algorithm based on heat transfer path decoupling (HPD-MODO) is proposed. The algorithm divides TMS into three heat transfer modes: oil-dominated, fuel-dominated and air-dominated, and adopts a case-partitioned strategy to acquire the optimal lubricating oil allocation among air-oil and fuel-oil heat exchangers. Across the three flight missions tested, HPD-MODO achieves solution accuracy equivalent to NSGA-II while reducing computation time from 2092~6173 ms to 1.36~1.51 ms. Based on proposed algorithm, heat sinks are reasonably allocated and the hot fuel return flow is minimized. In summary, the established strategy provides a flexible route to implement multivariable online optimization for intricate aircraft TMSs under next-generation flight platforms.

Understanding the rotational heat transfer characteristic of internal channels inside turbine blades depends on high-quality experimental research, and the excellent performance of the experimental apparatus is the primary factor in achieving this goal. This article introduces a new experimental apparatus. It adds a series of unique functions through significant structural innovation, improves the experimental ability of the original facility, and covers the working condition envelope of the current typical turbine blade internal cooling channel. The experimental apparatus can be divided into five subsystems. The air source and supply pipeline subsystem can realize dual separate coolant supply and combined coolant supply. The maximum rotational speed of the rotor subsystem can reach 1000 rpm, and the average rotation radius of the test section can be adjusted within the range of 0.27-0.75 m. Temperature and pressure measurement points have been increased to 256 and 8, respectively, with room for further upgrades. Moreover, the turntable design also provides a broad space for the application of advanced measuring instruments, which can meet the needs of the future such as infrared and liquid crystal temperature measurements. At present, the experimental facility has been put into use, and it has played an irreplaceable role in the rotational heat transfer research of turbine blade leading edge impingement cooling channel, double-wall cooling channel, and U-shaped channel with overflow under acceptable use costs.

The design of conformal intercoolers for advanced aeroengines is complicated by high compressibility and variable flow area. This paper presents a refined design methodology based on a fully coupled system of ordinary differential equations (ODEs), derived from first principles to solve axial gradients of static pressure and temperature. The model inherently captures friction, heat transfer, and area-change effects, overcoming limitations of conventional tools such as the LMTD method in converting between static and total parameters. The methodology is applied to design a conformal air-to-supercritical nitrogen intercooler and is rigorously validated against high-fidelity CFD simulations and a state-of-the-art segmented LMTD model. The results demonstrate that the proposed ODE model achieves excellent agreement with CFD in predicting both overall performance metrics and local parameter distributions. In contrast, the classic and enhanced segmented LMTD models exhibit significant deviations, with total pressure-drop prediction errors reaching 17.9-45.4 % on the air side, whereas the proposed ODE model reduces this to within 10 % under the design condition. These findings highlight the limitations of LMTD-based approaches in handling strongly coupled problems. This work provides a computationally efficient yet physically robust tool for the high-fidelity design and optimization of advanced thermal management components in high-speed propulsion systems.

Fin structures fabricated from Ceramic Matrix Composites (CMC) are increasingly employed in advanced aircraft cooling systems due to their low density, high thermal resistance and corrosion resistance. However, due to the inherent anisotropy and directionality in heat conduction, theoretical solutions for such CMC fin structures remain challenging, which limits their design and optimization for practical applications. Therefore, this paper proposes an analytical solution for the heat transfer rate and efficiency of anisotropic annular fins. A mathematical model was developed, yielding the governing dimensionless numbers based on the differential equations and boundary conditions. Solution methods included variable separation, Taylor expansion, and integral averaging. Derived analytical correlations were numerically verified. The numerical validation demonstrated that within the ranges of Bir = 0.005, Biz = 0.01 0.05, Biφ = 0.01 0.05, βR = 10, r* = 0.1, and α = 0 90°, the deviation in fin efficiency predicted by the proposed formula did not exceed 1.56

With the advancement of next-generation fighter aircraft, the escalating cooling demands of thermal management in aircraft and their engines are approaching the thresholds of conventional heat sinks, including ram air and fuel. A variable cycle engine (VCE), characterized by its third-stream design, facilitates potential multi-heat sink coordination within the fuel thermal management system (FTMS). Despite the use of decoupled VCE and FTMS modeling in previous research, the heat sink potential of internal secondary bypass air remains largely unexplored and unquantified, with its feedback effects on VCE energy efficiency also lacking rigorous investigation. Driven by the background, this study proposes a novel coupling of VCE and FTMS design. By leveraging multidisciplinary simulations, we provide the first quantitative analysis of the heat sink efficacy of secondary bypass air across representative flight missions and elucidate its synergistic mechanism with fuel. Investigations reveal that compared with ram air, secondary bypass air markedly reduces the thermal accumulation by 36.57%–74.06%. This improved thermal performance is accompanied by a 2.17%–4.10% decrease in the hot-return fuel flow. Intriguingly, the induced specific fuel consumption penalty throughout various typical flight missions consistently remains below 0.8%, thereby demonstrating the economic efficiency and sustainable benefits of employing secondary bypass air for thermal management. Furthermore, this study presents the first optimization strategy for allocating heat transfer areas. Specifically, an area ratio of 0.6 between the ram air and secondary bypass air significantly lowers the system hot-return fuel temperature by 2.68%. This work validates quantitative evidence for secondary bypass air–FTMS coupling and establishes a foundation for system-level thermal management schemes in advanced fighter aircraft and engine designs.

A practical intermediate-fidelity framework for predicting three-dimensional crossflow conjugate heat transfer (CHT) is still lacking under turbine-airfoil-relevant conditions, where large temperature differences, variable properties, moderate-Mach turbulent convection, viscous dissipation, and finite wall conduction act simultaneously. This study develops a decoupled kernel-superposition analytical-numerical framework for smooth flat-plate surrogates, in which one-dimensional streamwise operators on the hot and cool sides are iteratively coupled with a three-dimensional solid-conduction problem until interfacial temperature and heat flux converge. Both wall-temperature and wall-heat-flux routes are supported; for the wall-heat-flux route, a unified treatment is introduced for unheated-start and streamwise-varying thermal histories. Variable-property and compressibility/dissipation corrections are calibrated against two-dimensional boundary-layer RANS data over temperature ratios of 0.43-2.33, Reynolds numbers up to 3.6 & times;106, and Mach numbers up to 0.7. Validation against fully coupled three-dimensional crossflow CHT references over effective Biot numbers of 0.05-3.5 yields main-plate spanwise-averaged Nux errors of O(2%) and all-region errors within about 6%, while reducing total CPU time by more than two orders of magnitude over the present database. Proof-of-concept ribbed-wall and literature-derived film-cooling extensions further demonstrate that non-trivial streamwise boundary conditions can be incorporated within the same decoupled loop. The framework therefore provides an intermediate-fidelity tool between empirical correlations and full three-dimensional CHT for rapid thermal assessment, parametric studies, and design-oriented screening of turbine cooling configurations.

Accurate and computationally efficient prediction of thermal performance in serpentine tube heat exchangers (STHEs) is paramount for advanced aeroengine thermal management systems. Existing methods, however, often oversimplify complex flow phenomena, rendering them inadequate for identifying intricate local cross-flow interactions and their consequences, which directly impact system reliability and safety. To address this critical research gap, this paper introduces a novel analytical model for multi-pass STHEs that achieves both high accuracy and superior computational efficiency. Our model employs a unit-based approach, discretizing the STHE into repeatable single-tube row modules. By sequentially solving these modules, we achieve an accurate resolution of detailed, two-dimensional temperature distributions across the entire heat exchanger. A significant contribution of this work is the model's capability to predict and mechanistically explain the observed local inverse heat transfer phenomenon in parallel-flow configurations. This critical behavior, largely overlooked by conventional models, directly influences overall heat transfer performance. Rigorous numerical validation confirms the model's accuracy, with average deviations below 1% for the total heat transfer rate and 3.49% for the temperature distribution. Furthermore, comparative analyses reveal that our proposed analytical model significantly outperforms traditional methods in predictive accuracy. The improvement is larger than 5% when μ > 0.30 and NTU2 > 2.83 for parallel-flow STHEs, and μ < 0.43 and NTU2 > 0.56 counter-flow STHEs. These findings collectively demonstrate that the developed framework not only provides new physical insights into the heat transfer mechanisms governing multi-pass STHEs but also serves as an efficient computational tool for performance optimization of aerospace thermal management systems where stringent reliability constraints exist.

Research on turbine blade internal cooling channels is gradually transitioning from simplified structural channels to more intricate ones that closely approximate actual blade cooling channels. The present study conducts an experimental and numerical research to investigate the flow and heat transfer behaviors within a composite cooling channel under static and rotating states. The studied composite channel contains a U-shaped channel and a lateral outflow pin-fins channel, which are interlinked by middle axial jet holes. The inlet Reynolds number (Re), temperature ratio (TR), and rotation number (Ro) studied vary from 10,000 to 40,000, 0.07 to 0.14, and 0 to 0.33, respectively. Under static states, axial jet inflow along the pass can avoid the formation of low heat transfer zones within pin-fins channel. Axial jets impact on pin-fins can increase heat transfer by 17 % to 28 %. Under rotating states, the lateral outflow by axial jet holes can weaken rotating influence and reduce heat transfer discrepancy between the leading and trailing walls of the second pass. Along the second pass, the reduction in fluid velocity results in a trade-off between Coriolis force and buoyancy force, which has a significant effect on secondary flow, resulting in a higher heat transfer on the leading wall at TR = 0.07 than TR = 0.14, with a difference of up to 27 % in some zones. The axial inflow avoids the influence of Coriolis force, thereby ensuring consistency in heat transfer on the leading and trailing walls of pin-fins channel. Moreover, the flow resistance is also investigated.

To manage increasing thermal loads in aeroengines, this study proposes a novel surface air-oil heat exchanger (SAOHE) with microchannel fins that eliminates the need for high thermal conductivity materials. Compared to conventional titanium SAOHEs, the new design improves the comprehensive performance which defined as the ratio of heat transfer rate to pressure drop under high Reynolds number conditions (the ratio of the evaluation index Rp,f/Rp,mf exceeds 1). A new dual-working-fluid method is introduced to determine the airside heat transfer coefficient (HTC) without relying on accurate oil-side HTC measurements. Notably, when the thermal resistance ratio of the hot and cold sides exceeds 9 and 19, the deviation of the cold side thermal resistance calculated with and without accounting for hot side thermal resistance remains under 10% and 5%, respectively. Based on these findings, a new air-side Nusselt number correlation was developed. The experimental results significantly higher than the values predicted by Chang and Wang correlation for louvered fins and slightly lower than the values predicted by Gnielinski correlation, offering valuable guidance for the design of high-performance heat exchangers in aeroengine applications.

Non-uniform wall temperature or wall heat flux in internal convection produces streamwise non-local thermal-history effects that are not represented by local heat-transfer-coefficient closures calibrated under uniform wall conditions. Streamwise thermal-response superposition (TRS) is a compact reduced-operator strategy: it retains upstream thermal memory, accepts either temperature- or heat-flux-driven wall inputs, and enables rapid repeated response updates without resolving the full flow field. However, an internal-channel TRS formulation that couples Graetz-type thermal-entry response, bulk-temperature evolution, and local NuD prediction remains lacking. This work develops such a Graetz-type TRS model for stationary smooth square channels by transferring canonical thermal-entry responses through a hydraulic-diameter-scaled streamwise kernel representation. Two complementary routes are formulated: the T-route maps Tw(x) to wall heat flux, bulk temperature, local heat-transfer coefficient, and NuD, whereas the q-route maps qw(x) to wall temperature and the same output quantities. The model is assessed against square-channel CFD references using prescribed wall histories, one-sided active-wall tests, and a counterflow coupled-wall demonstration. For the prescribed-boundary and one-sided active-wall cases, post-cut RMS relative errors of NuD are 0.6–3.8%, while full-interval errors are 4.7–8.6% when the finite-mode-sensitive near-origin region is included. In the counterflow demonstration, both routes give post-cut NuD errors below 2.5% and reduce aggregate CPU time by factors of 5.30×10^3–7.70×10^3. The model is therefore positioned as a compact streamwise-causal internal-response closure for repeated thermal-response evaluation in smooth square-channel problems.

Constrained by the geometry of the turbine blade, traditional trailing edge cooling channels often adopt a flow pattern with bottom inlet and sidewall bleeding outlet. This leads to poor flow and consequently ineffective cooling in the channel upper region. Introducing a second inlet is an effective strategy to enhance cooling there. This study numerically investigates the flow heat transfer characteristics in a two-inlet rectangular pin-fin roughened channel under stationary and rotating conditions. The major inlet Reynolds number is 20000; the rotation number and inlet mass flow rate ratio (MR) range from 0 to 0.58 and 0 to 2. Numerical results are validated against experimental data. By measuring the flow distribution through sidewall bleeding slots, the experimental assumption of a linear streamwise decrease in local mass flow is corrected. Subsequently, the heat transfer data from different positions are unified using local MR. A critical local MR of 1 is observed, where two opposing fluids with equal flow rates meet, cancelling momentum and creating a low-velocity region with minimal heat transfer. Rotation shifts this weak region towards the low-radius area. There is also a critical local rotation number of approximately 2 in the one-inlet channel. Buoyancy force positively affects heat transfer, enhancing it by 45% in the one-inlet channel as the temperature ratio increases from 0.06 to 0.22. However, the enhancement reduces to 16% in the two-inlet channel. Local parameters provide detailed insights into the local flow heat transfer, offering a superior framework for the refined design of cooling channels.

This paper presents a study on constructing a digital twin system for an aircraft thermal management system (TMS) using a five-dimensional framework. The TMS is crucial for maintaining the operational efficiency and safety of aircraft by regulating onboard temperatures. Airborne temperature-dependent physical architecture provides thermal sensing, while the software system issues commands for coordinated regulation. The digital twin integrates physical and virtual entities through a user-friendly interface, enabling real-time monitoring and analysis of the system’s performance. The study focuses on a simplified TMS laboratory bench, comprising five modules: fuel supply, pressure boost, airplane heat generation, main combustion chamber flow control, and fuel scavenge. The digital twin includes a physical entity with sensors, a virtual entity modeled using Amesim software, digital data storage, and a connection dimension for data transmission. The system uses a PLC to handle data interaction and PID control algorithms to manage the equipment. Experiments were conducted to verify the digital twin’s accuracy. The results showed that the physical and virtual entities had high compatibility, with minimal differencesin parameters such as total mass flow, main combustion chamber mass flow, fuel scavenge mass flow, and temperature. This study demonstrates for the first time a system-level implementation of a five-dimensional digital twin framework on a simplified aircraft TMS, achieving less than 2

Modern advanced turbine blade mid-chord cooling systems typically have three passages with different geometric shapes and cooling schemes. The current study conducts experimental and numerical analysis of the aerothermodynamic performance in a blade-shaped serpentine channel. The channel features asymmetric cross sections, 180-degree tip and hub turns, a minor secondary inlet, staggered ribs and bleed holes. The main inlet Reynolds number (Re) and rotation number (Ro) respectively vary between 17000 and 33000 and from 0 to 0.4, and the mass flow ratio of the minor secondary coolant to the main (MR) ranges from 0 to 0.2. It is revealed that the flow interactions between bleed holes and ribs significantly improve wall heat transfer. The rotation effect on heat transfer is less pronounced in a realistic channel than in a smooth one. The minor secondary stream can increase the channel heat transfer, and the ideal MR falls between 0.1 and 0.15. The proportion of the mass flow rate of each bleed hole to the total remains almost consistent regardless of the Re and Ro. Finally, the correlations of averaged heat transfer with high accuracy (<= 10 %) are developed, which could interest turbine blade researchers and designers.

This study outlines the design and manufacture of three types of finned tube bundle heat exchangers (HX) for advanced aero-engines followed by a series of experimental evaluations on flow and heat transfer characteristics. Considering the compactness and reliability, the finned tube bundle HX is more suitable for aero-engines when the heat transfer unit is small-diameter serpentine tubes with the enhanced heat transfer area provided by fins. In this paper, the Logarithmic Mean Temperature Difference method (LMTD) was utilized for the design of a small- diameter (OD: 3.6 mm with 0.3 mm thickness) finned tube bundle HX. Mass reduction was achieved by perforating the fins and removing the support devices, respectively. The flow and heat transfer characteristics of HXs and the impact of the mass reduction schemes were evaluated based on a series of comparative experiments. Among them, the connectionless scheme can significantly improve the power-to-mass ratio, providing it certain application value despite causing considerable flow resistance. Subsequently, empirical correlations for the flow and heat transfer outside the tubes suitable for small-diameter finned tube bundle HXs were proposed based on the experimental results, with over 95% of the data falling within a 10% error margin. These correlations have substantially modified the previous one for large-diameter straight tubes, providing an important reference for the future design of HXs in aero-engine.

In line with increases in aircraft engines’ performance, cooled cooling air technology (CCA technology) has been put forward to address thermal management issues. It mainly utilizes bypass air to cool the compressor bleed air for turbine cooling by the application of a heat exchanger. However, the current serpentine-tube CCA heat exchanger results in significant pressure losses, which have a detrimental effect on both the engine’s total thrust and fuel consumption. Therefore, two novel heat exchangers with ultralow flow resistance characteristics are proposed, and their thermal and hydraulic performances are experimentally tested. Besides, a multidimensional evaluation method is established by simulation work which integrates the heat exchanger itself with the whole engine system performances (including total thrust and specific fuel consumption). In regard to versatility, this work also explores the superiority of these novel types of heat exchangers in precooling technology. Our findings prove that the two novel heat exchangers take an over 50 % lower pressure drop compared to the traditional serpentine-tube type under the same heat transfer conditions. When they are applied to CCA or precooling technology, the enhancement in total thrust FN and reduction in specific fuel consumption sfc have been observed under different flight missions. When using the unilateral plate-fin heat exchanger, the maximum increment in total thrust is approximately 51.5 %, and the decrease in sfc is approximately 33.4 %. And those for the longitudinal flow tube bundle heat exchanger are 38.1 % and 27.2 % respectively. Further analyses reveal that the superiority is mainly caused by the increase in mass flow rate and exhaust gas pressure. To sum up, the present study provides a theoretical and experimental basis for the design of advanced heat exchangers in the aerospace field, and further optimization work is still needed to enhance their heat transfer performance.

High thermal loads in aviation necessitate compact, efficient, small-diameter tube bundle heat exchangers. However, the combined impact of high velocity and small diameter remains unclear. To address this gap, the present study investigates the airside thermal-hydraulic performance of small-diameter tube bundles in compressible cross-flow. Contrary to classical correlations, present experiments measuring Nusselt number (Nu) and friction factor (fac) reveal significant diameter dependence at high Reynolds numbers (Re). At Re = 10,000, 1 mm tube bundle exhibits an increase of 25.2 % in Nu and 17.5 % in fac compared with its 5 mm counterpart. To model tube conjugate heat transfer, a refined dimensionless parameter framework is developed to comprehensively incorporate the complex interactions among Re, Prandtl number (Pr), Mach number (Ma), and Eckert number (Ec). Numerical simulations reveal that the coupling between Ma and Ec plays a pivotal role in determining flow behavior and heat transfer characteristics. Specifically, higher Ma promotes flow separation, increasing pressure drop alongside enhanced heat transfer. In contrast, higher Ec suppresses flow separation, reducing pressure drop while still augmenting heat transfer. These findings elucidate the intricate interplay of thermal-hydraulic mechanisms in small-diameter tube bundles under compressible flow, providing a fundamental basis for designing advanced aerospace compact heat exchangers.

The longitudinal compact arrangement (LCA) tube bundle shows great potential for advanced compact heat exchangers, with its excellent properties of low flow resistance and compactness. However, the lack of specific performance data and reliable empirical correlations hinders its practical application. This study experimentally investigates the airside heat transfer (Nusselt number, Nu) and flow resistance (friction factor, f) characteristics of LCA tube bundles across a wide range of transversal pitch ratios (S1/D = 1.521 to 8.621) for Reynolds numbers (Re) from 1000 to 10,000. The novelty of this study lies in developing the first comprehensive set of empirical correlations specifically for LCA tube bundles; these correlations are validated against experimental data with predicted deviations under 10 % for Nu and 15 % for f. Key findings reveal that while LCA tube bundles exhibit moderately lower Nu compared to typical inline tube bundles, they achieve significantly lower friction factors. Consequently, a comparative analysis using the Colburn j-factor to friction factor ratio (j/f) shows the overall thermal-hydraulic performance (j/f) of the LCA tube bundle is 1.2 to 2.65 times greater than that of typical inline tube bundles, primarily attributed to its ultralow flow resistance characteristics. This research provides crucial, validated design tools and quantitatively demonstrates the superiority of the LCA tube bundles for applications prioritizing low pressure drop alongside effective heat transfer, facilitating the design of ultralow flow resistance compact heat exchangers.