The trailing edge of the high pressure turbine blade presents significant challenges to the turbine cooling engineer. A novel cooling design using cross corrugated slots for the trailing edge has been proposed. This geometry allows blade designers to finely tune pressure loss and consequently coolant flow through the slot, but potentially results in poor film cooling performance downstream of the slot exit, an effect that could be mitigated with exit shaping. The current study is focused on comparing film cooling effectiveness on the cutback surface and lands with a plain rectangular slot under the same conditions. A set of nine cross corrugated internal slot geometries has been investigated in a large scale model of the trailing edge pressure side ejection slot exit. Four geometries used a 90° included angle with variations to the channel alignment at slot exit. Four used a 120° included angle, with the same variations to the exit alignment. The final geometry used a 90° included angle with exit shaping. Pressure sensitive paint was used to measure adiabatic film cooling effectiveness at five blowing ratios ranging from 0.6 to 1.4 in increments of 0.2. High resolution 2D distributions of film cooling effectiveness both on the cutback surface and the top of the lands were recorded. It was found that unmodified cross corrugated slots do result in poor film effectiveness on the cutback surface compared to a plain rectangular slot. However, land cooling is slightly improved, and applying exit shaping to the cross corrugated slot results in effectiveness levels at the trailing edge on par with or even superior to the rectangular slot at blowing ratios of 0.8 or below. Therefore, in this respect, the novel cross corrugated slot design proposed is a viable candidate for blade design, provided exit shaping is used and low blowing ratios are expected.

The trailing edge of the high pressure turbine blade and vane presents significant challenges to the turbine cooling engineer. The current research has focused specifically on the effect of cutback surface protuberance, or "land", shapes on film cooling effectiveness. A set of six different land geometries has been investigated in a large scale model of the trailing edge pressure side ejection slot exit. Slot height and width and lip height was maintained. Pressure sensitive paint was used to measure adiabatic film cooling effectiveness at five blowing ratios ranging from 0.6 to 1.4 in increments of 0.2. High-resolution full surface distributions of film cooling effectiveness both on the cutback surface and the top of the lands were recorded. It was found that tapering the lands did not significantly increase effectiveness on the lands and slightly reduced effectiveness near the lands. Using a diffuser shape improved average effectiveness greatly and gave the best overall performance up to the end of the lands except at the lowest blowing ratio of 0.6, where having no lands was slightly better.

This research focuses on film cooling of the trailing edge of a scaled up turbine rotor blade with engine-representative Mach number distribution. Pressure sensitive paint was used to obtain high-resolution adiabatic film cooling effectiveness measurements in the trailing edge region of the scaled turbine blade. The large scale, high-speed experimental set-up consists of a Perspex test section for maximum visibility of the PSP coated blade. The test section was designed to recreate a single blade passage of a gas turbine with inlet Mach and Reynolds numbers matching the corresponding values in an engine. The test blade has a constant cross section, representative of the mid-span profile of the high pressure turbine rotor blade. It was manufactured from aluminium to minimize temperature gradients over the surface of the test blade. In the current research, pressure surface cooling slots at the trailing edge were examined and the effect of cutback surface protuberance, or ‘land’, shapes on trailing edge film cooling was studied. Nitrogen and air were used as coolant gases giving a coolant to mainstream density ratio close to 1. Two land geometries-straight and tapered-were studied for a set of 6 blowing ratios from 0.4 to 1.4 in steps of 0.2. Land taper has a benefit for film cooling near the slot exit but its advantage reduces close to the trailing edge. For both geometries, film effectiveness falls with blowing ratio from 0.4 to 0.8 and increases with blowing ratio in the 0.8 to 1.4 range. Crossflow causes the coolant film to be biased towards one side of the lands. Film effectiveness results are compared with data from a scaled up low speed flat plat model of the trailing edge to explain the effect of acceleration on film cooling.

Adiabatic film cooling effectiveness measurements are made on nozzle guide vane leading edges in an engine-realistic flow environment. The tested leading edges feature radial showerheads with different spanwise distributions of hole surface angle. The showerheads blow towards the midspan, except for one model with showerhead holes orthogonal to the vane surface. The results show that low surface angle radial showerhead holes generate high effectiveness within their rows and further downstream, but neglect the stagnation region lying between the two most upstream cooling hole rows. This downstream effectiveness gain is due to both the continued surface attachment of this coolant as it progresses downstream, and its beneficial interactions with downstream cooling jets. Moderate radial showerhead surface angles cause moderate coolant jet penetration into the mainstream, which promotes near-surface mixing of the coolant with the mainstream, increasing stagnation region effectiveness. The mixing effect is enhanced by the intense turbulence generated by combustor dilution jets. High surface angles may cause the stagnation region coolant to penetrate too far for either of these gains to be realised. Considering also the presence of endwall film cooling, these effects, taken together, suggest the superiority of radial showerheads which blow towards the midspan, as against those which blow towards each endwall. Surface temperature data is acquired by a novel infrared thermography technique which permits measurement of both heat transfer coefficient and film effectiveness from a single heated test.

The trailing edge of the high pressure turbine blade and vane presents significant challenges to the turbine cooling engineer. The cooling design using cross corrugated slots allows for tuneable coolant flow through the slot, but results in poor film cooling performance, an effect that must be mitigated with exit shaping. A set of four cross corrugated slot geometries with optimised exit shaping has been investigated in a large scale model of the trailing edge slot. Pressure sensitive paint was used to measure adiabatic film cooling effectiveness at blowing ratios ranging from 0.6 to 1.4. The new optimised designs were able to improve upon previous designs, permitting levels of film effectiveness on the cutback surface up to the end of the lands that were significantly better than a plain rectangular slot at blowing ratios of 0.8 and below, while maintaining the pressure loss benefits of the cross corrugated slot and also increasing effectiveness next to the lands.