MUMBAI, India, Sept. 28 -- Intellectual Property India has published a patent application (202641112666 A) filed by G Ashwin Prabhu; Dr. G. Ashwin Prabhu; Mr. Mughilan S; Mr. Veenish A; Mr. Sai Vengatesh S; and Sasidharan R on September 20, 2026, for Investigation And Design Of Advanced Hybrid Thermal Barrier Coated And Multi-Pass Film-Cooled High-Pressure Turbine Blades For Jet Engine Propulsion Systems.
Inventors include Dr. G. Ashwin Prabhu; Mr. Mughilan S; Mr. Veenish A; Mr. Sai Vengatesh S; and Sasidharan R.
The application for the patent was published on September 25, 2026, under issue no. 39/2026.
Abstract: High-pressure turbine (HPT) blades in modern aviation gas turbine and jet engines operate under extreme thermo-mechanical environments characterized by severe thermal gradients, high rotational stresses, aerodynamic shear, and aggressive oxidative and hot-corrosive atmospheres. The present invention relates to an advanced high-temperature turbine blade assembly and structural cooling architecture for aerospace jet engines. The blade assembly comprises a single-crystal (SX) nickel-base superalloy core integrated with a multi-layered functionally graded Thermal Barrier Coating (TBC) system and an internal-external hybrid cooling network. The functional coating stack comprises a metallic MCrAlY (where M = Ni, Co) bond coat, an intermediate thermally grown oxide (TGO) suppression layer, and a columnar-structured rare-earth yttria-stabilized zirconia (7-8 YSZ) ceramic topcoat applied via Electron Beam Physical Vapor Deposition (EB-PVD). The internal structure incorporates an aerodynamically contoured 3-pass serpentine cooling passage equipped with optimized 45° staggered rib turbulators, leading-edge impingement cavities, and trailing-edge pin-fin pedestal arrays. Cooling air bled from the high-pressure compressor stage traverses the serpentine passage and discharges through an array of compound-angled, shaped diffusion film-cooling micro-orifices positioned along the leading edge, pressure side, and suction side. The interaction between the functionally graded TBC architecture and the active boundary-layer film cooling provides a thermal drop across the wall thickness of up to 280 °C under continuous turbine entry temperatures (TET) exceeding 1650 °C. Microstructural analysis and computational thermal-fluid evaluations indicate enhanced resistance to thermal barrier spallation, reduction in thermal fatigue cracking, prevention of film-coolant jet blow-off, and substantial suppression of microstructural creep deformation. The proposed architecture significantly enhances the thermodynamic cycle efficiency, structural durability, and operational lifecycle of critical jet engine components.
Disclaimer: Curated by HT Syndication.
