To determine the extent of overheating/burning, microstructural studies and hardness measurements were made along the nozzle body from the stem end to the flange end. Figure CH26.2 shows a sketch of the nozzle body delineating zones of different micro-structures and hardness. Analysis of the nozzle material in SEM indicated that it was made of chromium steel. The other atomizing parts found inside the nozzle had not suffered any damage due to heat. The Widmansta¨tten/basket weave microstructure seen near the flange and the stem is indicative of overheating above the trans-formation temperature (950⬚C, or 1740 ⬚F) followed by relatively fast cooling. This is also corroborated by the variation in hardness on the stem. The overheating was confined mostly to the two ends and had not affected the entire stem of the nozzle. This suggests that the stem had not been exposed to high temperature to the same extent as the ends. Thus, it can be concluded that the exposure to high temperature is very short; a “guess estimate” may be of the order of minutes.
Fig. CH26.1 Damaged fuel nozzle
DOI:10.1361/faes2005p126 www.asminternational.org
Discussion
The damages observed on the nozzle were due to wrong mount-ing. By that, the primary and secondary manifolds also get
inter-changed. When the nozzle is mounted with the atomizer toward the compressor side, the fuel spray will be turned back to envelope this end over the stem. There is no way of igniting the fuel at this point. When the heat radiated from the turbine heats the diffuser parts to match the flash point of the fuel, it can become ignited, stabilizing a flame around the stem causing it to burn. This flame might be responsible for the overheating/melting of some areas on the nozzle.
Conclusion
The variation in the microstructure on the stem of the nozzle indicated that the nozzle had been subjected to very high tem-perature close to the melting point of the nozzle material for a short duration. This was due to the reverse mounting of the nozzle in the combustor.
Recommendation
The end manifolds should be designed in such a way that the fuel nozzle can be fitted facing the turbine only.
Fig. CH26.2 Sketch of the nozzle assembly showing zone 1, Widman-sta¨tten/basket weave microstructure (hardness, 320 HV);
zone 2, spheroidized microstructure (hardness, 410 HV); and zone 3, marten-sitic microstructure (hardness, 520 HV)
CASE 27
Failure of a First-Stage
Compressor Blade in an Aircraft Engine
Summary
The first-stage compressor blade of an aircraft engine failed.
Investigations revealed that the failure was due to fatigue, the fa-tigue crack originating from corrosion pits at the root transition region on the convex side of the airfoil.
Background
A blade on the first stage of the compressor of an aircraft engine failed.
Pertinent Specifications
The blade was forged out of U961W material and heat treated as per schedule. It was Ni-Cd plated on the airfoil, silver plated on the dovetail area, and chromium plated on the bottom of the blade dovetail.
Visual Examination of General Physical Features
Figure CH27.1 shows the failed blade. It had failed at the root in the region of transition from the dovetail to the blade airfoil.
The fracture surface showed two distinct regions: a smooth region extending inward from the surface to a depth of 50% of the cross section and a rough region extending over the remaining cross section at the root (Fig. CH27.2).
Testing Procedure and Results
Macroscopy
Under the stereobinocular microscope, a series of half-moon-shaped beach marks were seen in the smooth region (Fig. CH27.3).
These marks are typical of fatigue crack propagation. The inner-most mark was found to surround the midpoint of the root tran-sition region on the convex side of the blade airfoil, indicating this to be the point of crack initiation. The width of the fatigue crack was 75 mm. The remaining cross section of the blade at the root was rough, typical of single overload fracture. At the point of crack initiation, presence of a corrosion pit was revealed (Fig. CH27.4).
Pitting was also seen in the surrounding region including the root flank and the root landing where the blade fits into the dovetail.
Chemical Analysis
EDAX analysis in SEM showed that the blade was made of 12% Cr, 1.2% Ni steel.
25 mm
Fig. CH27.1 Failed first-stage compressor blade
4 mm
Fig. CH27.2 Fracture surface of the blade showing two distinct regions
DOI:10.1361/faes2005p128 www.asminternational.org
Scanning Electron Fractography
In the SEM, poorly defined markings resembling fatigue stria-tions were seen in the half-moon-shaped regions of the fracture surface. Corrosion pits were also seen at the point of crack origin and its surroundings. No foreign object damage or material flaw was seen.
Discussion
The presence of well-defined beach marks and their extent on the fracture surface indicates that the blade failed by a fatigue crack initiating at the root transition region on the convex side of the
airfoil and propagating inward to a depth of 50% of the blade root cross section before giving rise to a single overload failure. The crack initiation can be traced to the corrosion pits. No tool mark or foreign object damage or flaw in the material is responsible for the initiation of the fatigue crack.
Conclusion
The compressor blade failed by a fatigue crack initiating from corrosion pits at the root transition region on the convex side of the airfoil and propagating to a depth of 50% of the blade cross section, before the final overload fracture. No foreign object dam-age or material flaw was responsible for the initiation of the fatigue crack.
Fig. CH27.3 Beach marks in the smooth region, typical of fatigue
1 mm
Fig. CH27.4 Corrosion pit seen at the region of fracture origin