**2. Experimental setups and test facilities**

The Institute for Thermal Turbomachinery and Machine Dynamics at Graz University of Technology operates a 3 MW compressor station in order to supply a couple of test facilities continuously with pressurized air. Among them are the subsonic test turbine facility for aerodynamic, acoustic, and aeroelastic investigations (STTF‐AAAI) as well as the two‐spool Transonic Test Turbine Facility (TTTF).

#### **2.1. Subsonic test turbine facility for aerodynamic, acoustic, and aeroelastic investigations**

In the described subsonic turbine test facility for aerodynamic, acoustic and aeroelastic investigations (STTF‐AAAI), the maximum pressure ratio is limited to 2 due to the inlet spiral casing. The mass flow rate is limited to 15 kg/s due to the compressor station characteristic. A temperature at stage inlet of max. 100°C can be realised. This inlet temperature can be adjusted within a wide range by means of water air coolers. The pressurized air enters the facility through a spiral inlet casing where the flow is turned into axial direction. That spiral inlet casing also supports the front bearing of the overhung‐type turbine shaft. The shaft is coupled to a water brake counteracting the power of the turbine. The necessary cooling water cycle of the brake is connected to the re‐cooling plant of the institute.

For a test rig it is mandatory to provide well‐defined and uniform inflow conditions; therefore a de‐swirler and a perforated plate are located far upstream of the inlet guide vanes. That mentioned inlet guide vanes upstream of the stage (and downstream of the perforated plate) simulate additional wakes of other upstream low pressure turbine stages. The air leaves the test rig through a cylindrical acoustic measurement section, supporting struts centring the acoustic measurement section, exhaust casing, and the exhaust stack to ambient. For more information a detailed description of the subsonic test turbine facility can be found in Ref. [1].

#### *2.1.1. Turbine stage and turbine exit casings[SEQA1]*

At the beginning of the jet era noise emissions were dominated by jet mixing noise. That has beenreducedwhenintroducingmodernbypassenginesbyloweringthejetspeed.Noiseemitted by fan, turbine, and compressor has then become important with the reduction in jet speed. Now,thatnoisehas tobe reducedsignificantly.Since thepublicationoftheACAREgoals,which are targets defined by the Advisory Council for Aeronautics Research in Europe to reduce the environmental impact of air transportation, the commercial and political pressure to reduce CO2, NOx and noise (up to 20 dB reduction of noise level until 2020 compared to technologies of the year 2000) has been increased considerably. A lot of research over the past decades was done reducing fan and compressor noise in orderto achieve the noise reduction targets. During the last years, noise emission from the fan was much reduced that suddenly noise from the interactionofthe last stage lowpressure turbine andthe turbine exitguidevaneofthe exit casing became perceivable. Nowadays also manufacturers of low pressure turbine components have toconsideracousticaspects intheirdesigntobeabletoreachtheACAREgoals in2020.Therefore a lot of research is currently done in that field of expertise. An additional benefit of that low noise levels is that passengers as well as residents living in the vicinity of airports feel more comfortable. Basically an increasing acceptance to live close to airports and rise of life quality

can be achieved if the noise level of aero engines is decreased significantly.

In this chapter measures to reduce noise generated and propagating from modern aero engine turbines are presented. The main issue is to find methods and/or new engine designs that reduce noise without causing considerable losses or a reduction of thrust. However, a lot of novel engine architectures are investigated in several national and international funded projects with the goal to reduce the emission of pollutants, e.g. by lowering the engine weight. This can be achieved by reducing the length of the entire engine by reducing the axial spacing between blade rows or integrating additional functions in one part, e.g. a non‐lifting strut in an intermediate turbine duct which also has to turn the flow and provide the next rotor with the correct inflow conditions. This leads to a so‐called turning mid turbine frame. However, all modifications on these parts of the engine will influence the noise generation and propa‐

gation. Also a considerable change of excitation of blades and vanes can be observed.

The Institute for Thermal Turbomachinery and Machine Dynamics at Graz University of Technology operates a 3 MW compressor station in order to supply a couple of test facilities continuously with pressurized air. Among them are the subsonic test turbine facility for aerodynamic, acoustic, and aeroelastic investigations (STTF‐AAAI) as well as the two‐spool

**2.1. Subsonic test turbine facility for aerodynamic, acoustic, and aeroelastic investigations**

In the described subsonic turbine test facility for aerodynamic, acoustic and aeroelastic investigations (STTF‐AAAI), the maximum pressure ratio is limited to 2 due to the inlet spiral casing. The mass flow rate is limited to 15 kg/s due to the compressor station characteristic. A

**2. Experimental setups and test facilities**

Transonic Test Turbine Facility (TTTF).

4 Recent Progress in Some Aircraft Technologies

During the aerodynamic design process it was crucial to achieve relevant model parameters to reproduce the full scale low pressure turbine (LPT) configuration. The diameter of the test rig rotor is about half of that of a commercial aero engine LPT. Therefore the rig is operated at higher rotational speeds to realise an engine relevant loading parameter. A meridional section of the rig is shown in **Figure 1**. The state‐of‐the‐art (reference) and the leaned turbine exit casing (TEC) are shown in the sketch at the top (a). The inverse cutoff as well as the high lift design (H‐TEC) can be seen at the bottom (b). It has to be mentioned that the bladings are not drawn to scale. The rig is characterised by a high aspect ratio unshrouded low pressure turbine rotor followed by the TEC. Relevant geometry parameters can be seen in the upper half of **Table 1**.

**Figure 1.** Meridional section of the STTF‐AAAI; (a) reference and leaned TEC and (b) inverse cutoff and H‐TEC.

Four different TEC setups with different vane counts (see **Table 1**) have been tested but the leading edge is at the same axial position for all configurations. One significant difference is that the reference and leaned TEC are manufactured without fillets while the TECs with smaller chord length have fillets at hub and tip due to manufacturing and assembly require‐ ments. The leaned TEC was optimised (detailed information can be found in Ref. [2]) in order to reduce rotor‐TEC interaction noise by keeping the profiles of the turbine exit guide vanes (TEGV) to be able to lead through the same supply lines as through the reference TEC. As it was shown in some European projects, e.g. DREAM, the rear bearing can move forward under the TCF section for future engine architectures giving the designer the opportunity to aero‐ dynamically and/or acoustically optimise the vanes of the turbine exit casing. Therefore, the third setup is an acoustically optimised TEC named inverse cutoff TEC. The basic idea of that setup is to utilise a small cutoff corridor in between two cuton regions. A detailed description can be found in Ref. [3] and a verification and comparison with experimental results is given in Ref. [4]. The fourth setup is aerodynamically optimised and is designed to reduce losses at aero design point.

Further, the rig has some inlet guide vanes (IGV) in order to impose some typical pre‐swirl on the flow. Stator vanes are located downstream of the IGVs, followed by the rotor and the turbine exit guide vanes (TEGV). The 1 and 1/2 stage is representative of the last stage of a commercial engine with TEGV. **Table 1** shows the blade count and the main geometrical details of the turbine. The rig in its current setup is characterized by a high aspect ratio unshrouded rotor followed by one of the above described turbine exit casings. The tip leakage flow dominates the flow field downstream of the rotor. The flow through the guide vanes is mainly influenced by secondary flows. The TEGV are designed to turn the swirling flow into an axial direction (reducing swirl and lower the kinetic energy of the flow) and to recover some static pressure.

Additionally, for this test rig, a second stator and low pressure turbine rotor has been designed. Stator and rotor have the same blade count as the reference design, but different profile geometries, including a rotor with a 20% increased loading parameter. The design intent of that second stage was to provide a similar/identical rotor exit flow as well as shaft power of the test rig. Because of the larger turning of that highly loaded rotor, the operating points must have been adjusted. A lower rotational speed in order to keep the power output identical was chosen. However, the stage pressure ratio has been kept the same as for the datum stage. There have been two reasons to keep the blade and vane counts identical. Firstly, it is the geometrical limitation of the test rig. Axial chord of both stator and rotor had to be the same as well as the axial distance between the vanes and blades. Secondly, the resulting acoustic and aerodynamic field is dependent on the number of blades and vanes of the test rig [5]. In order to keep the modal structure of the blade/vane interactions the same as for the datum stage, the number of blades and must be the same.

#### *2.1.2. Operating conditions*

Because the STTF‐AAAI is used for both acoustic and aerodynamic investigations, the main operating points are selected according to relevant noise certification points. They have been defined using a typical aerodynamic design point of a last stage of an LPT. That design point is derived from current LPT design practice and scaled along reduced speed, reduced mass flow (both referred to 288.15 K and 1013.25 mbar) and pressure ratio. For this investigation the acoustically relevant operating point approach was chosen. The Reynolds number of the TEGVs is defined using the midspan conditions at rotor exit as well as the axial chord of the vanes. The lower half of **Table 1** shows the operating conditions.


**Table 1.** Geometry details and operating conditions.

#### **2.2. Transonic test turbine facility**

Four different TEC setups with different vane counts (see **Table 1**) have been tested but the leading edge is at the same axial position for all configurations. One significant difference is that the reference and leaned TEC are manufactured without fillets while the TECs with smaller chord length have fillets at hub and tip due to manufacturing and assembly require‐ ments. The leaned TEC was optimised (detailed information can be found in Ref. [2]) in order to reduce rotor‐TEC interaction noise by keeping the profiles of the turbine exit guide vanes (TEGV) to be able to lead through the same supply lines as through the reference TEC. As it was shown in some European projects, e.g. DREAM, the rear bearing can move forward under the TCF section for future engine architectures giving the designer the opportunity to aero‐ dynamically and/or acoustically optimise the vanes of the turbine exit casing. Therefore, the third setup is an acoustically optimised TEC named inverse cutoff TEC. The basic idea of that setup is to utilise a small cutoff corridor in between two cuton regions. A detailed description can be found in Ref. [3] and a verification and comparison with experimental results is given in Ref. [4]. The fourth setup is aerodynamically optimised and is designed to reduce losses at

Further, the rig has some inlet guide vanes (IGV) in order to impose some typical pre‐swirl on the flow. Stator vanes are located downstream of the IGVs, followed by the rotor and the turbine exit guide vanes (TEGV). The 1 and 1/2 stage is representative of the last stage of a commercial engine with TEGV. **Table 1** shows the blade count and the main geometrical details of the turbine. The rig in its current setup is characterized by a high aspect ratio unshrouded rotor followed by one of the above described turbine exit casings. The tip leakage flow dominates the flow field downstream of the rotor. The flow through the guide vanes is mainly influenced by secondary flows. The TEGV are designed to turn the swirling flow into an axial direction (reducing swirl and lower the kinetic energy of the flow) and to recover some static

Additionally, for this test rig, a second stator and low pressure turbine rotor has been designed. Stator and rotor have the same blade count as the reference design, but different profile geometries, including a rotor with a 20% increased loading parameter. The design intent of that second stage was to provide a similar/identical rotor exit flow as well as shaft power of the test rig. Because of the larger turning of that highly loaded rotor, the operating points must have been adjusted. A lower rotational speed in order to keep the power output identical was chosen. However, the stage pressure ratio has been kept the same as for the datum stage. There have been two reasons to keep the blade and vane counts identical. Firstly, it is the geometrical limitation of the test rig. Axial chord of both stator and rotor had to be the same as well as the axial distance between the vanes and blades. Secondly, the resulting acoustic and aerodynamic field is dependent on the number of blades and vanes of the test rig [5]. In order to keep the modal structure of the blade/vane interactions the same as for the datum stage, the number of

Because the STTF‐AAAI is used for both acoustic and aerodynamic investigations, the main operating points are selected according to relevant noise certification points. They have been

aero design point.

6 Recent Progress in Some Aircraft Technologies

pressure.

blades and must be the same.

*2.1.2. Operating conditions*

The setups being tested consist of a single‐stage unshrouded transonic HP turbine. Down‐ stream of that turbine stage, a S‐shaped turbine centre frame, which is the main part of interest for this investigation, is located. The centre frame is followed by a shrouded counter‐rotating LPT rotor (see **Figure 2**). The pressurized air flows through the transonic unshrouded HPT rotor and enters the turbine centre frame. The air is then turned by 16 struts in the turbine centre frame (resulting in a turning mid turbine frame [TMTF]) in negative direction relative to the rotation of the HP rotor. After that the air enters the LP rotor at a larger diameter and with an appropriate swirl angle. Blading parameters and operating conditions can be seen in **Table 2**.

**Figure 2.** Meridional section of the TTTF.


**Table 2.** Blading parameters and operating conditions.

Configuration 1 (C1) consists of 16 turning struts. It has a non‐dimensional length of about 3.5 (Lax/hin) and an area ratio of 2. C1 was designed using a quite complex three‐dimensional design of the strut and keeping rotationally symmetric endwall contours. The struts have a maximum thickness to chord ratio of 22% at about 25% of the axial chord length to provide enough space for service lines like oil pipes and for load carrying structures. Two major goals were set for the aero design of the second configuration (C2). Firstly, C2 has to be more aggressive than the first configuration and secondly, the LPT performance has to be un‐ changed, keeping the same pressure loss. For C2 also, 16 turning struts have been assembled. For all configurations the same high pressure stage and low pressure rotor have been used. Therefore, the radial offset and the area ratio are the same for both setups but configuration C2 was designed to be 10% shorter than configuration C1. C2 has a non‐dimensional length of 3.1. Close to the hub the axial gap between the strut trailing edge and the leading edge of the LPT is the same as for C1. However, that gap is 20% shorter at midspan and 50% at the casing, respectively. The vane is designed to produce minimum blockage and therefore, it should create minimum losses. Both struts give the possibility to lead through identical service lines. To avoid additional losses and provide same inflow conditions to the LP turbine as the configuration C1, non‐axisymmetric endwall contouring was applied at the hub. The optimi‐ sation of endwall contour was performed using parameterization based on orthogonal basis perturbation functions. The third setup is based on the same geometry as C1, but 32 splitter vanes have been added for setup C3. Within each strut passage two non‐lifting splitters are located. The splitter vanes have been numerically designed and represent a compromise between aerodynamic effectiveness and additional blockage, because it was required to use the same casing parts. Therefore, it was not possible to change the area to count for the additional blockage of the splitters.

#### **2.3. Instrumentation**

**Figure 2.** Meridional section of the TTTF.

8 Recent Progress in Some Aircraft Technologies

**HP vane HP blade Strut**

Vane/blade no. 24 36 16 32 72 *h* / *cax* 1.15 1.37 0.46/0.53 3.5 2.94

*BPFHP kHz* 6.69 6.69 6.69

*BPFLP kHz* 4.26 4.26 4.26

Stage *pt* ratio HPT/ LPT 3/1.3 3/1.3 2.83/1.36 Power [MW] HPT/ LPT 1.44/ 0.3 1.44/ 0.3 1.43/ 0.3

Tip gap - Unshrouded - - Shrouded

Configuration 1 (C1) consists of 16 turning struts. It has a non‐dimensional length of about 3.5 (Lax/hin) and an area ratio of 2. C1 was designed using a quite complex three‐dimensional design of the strut and keeping rotationally symmetric endwall contours. The struts have a maximum thickness to chord ratio of 22% at about 25% of the axial chord length to provide enough space for service lines like oil pipes and for load carrying structures. Two major goals were set for the aero design of the second configuration (C2). Firstly, C2 has to be more aggressive than the first configuration and secondly, the LPT performance has to be un‐ changed, keeping the same pressure loss. For C2 also, 16 turning struts have been assembled. For all configurations the same high pressure stage and low pressure rotor have been used.

C1 C2 C3

**C1/C2/C3**

**Splitter C3**

**LP blade**

**Blading parameters**

Operating conditions ADP

**Table 2.** Blading parameters and operating conditions.

Five‐hole probes with a probe head of 2.5 mm diameter have been applied. The probes were used in measurement plane C downstream of the rotor and in plane D downstream of the TEGV (see **Figure 1b**). The probes are calibrated in a Mach numbers range between 0.1 and 0.8 in 0.1 steps, yaw angles between –20°and +20° in steps of 4 deg, pitch angles between – 16°and +20°, also in steps of 4°. Negative values of the yaw angle indicate a counter‐rotating flow and negative values of the pitch angle indicate the flow direction towards the hub. A multi‐parameter approximation correlates the calibration characteristics and the flow value to be measured.

The axial positions of measurement planes (marked with letters) can be seen in **Figure 1** for the STTF‐AAAI and **Figure 2** for the TTTF. TEGV/Strut inlet plane is located downstream of the rotor trailing edge. TEGV exit plane can be found 55% of the axial chord length of the H‐ TEC TEGV downstream of its trailing edge and also 130% axial chord length of the reference TEC downstream. The measurement grid covers one TEGV or a strut pitch and about 95% passage height. The five‐hole probe was traversed along radial lines. In each measurement point the probe has been aligned with the flow vector to reach the highest accuracy and further to ensure not to exceed the calibration range of the probe (with these probes, it would not have been necessary if one can ensure to be always within the calibration range).

In order to calculate the sound power (propagating downstream) from measured sound pressures, 12 flush mounted condenser microphones (1/4") at the hub and 12 at the casing have been applied. The microphones have been located in the rotatable acoustic measurement section that can be rotated 360° with arbitrary step size. In addition to these microphones one additional microphone was mounted at a stationary position downstream of the TEGV's as well as struts trailing edge and is used as a reference. The complete sound field was detected at the hub and the casing by traversing the section 360° in steps of 2°. Some more detailed information about the acoustic measurement section is given in Moser et al. [6] and in Faustmann et al. [7].

#### *2.3.1. Measurement uncertainty*

The measurement system is made up by 11 multichannel pressure transducers PSI 9016 with a total amount of 176 channels and an accuracy of 0.05% full scale for pressure measurements. Four National Instruments Field Point FP‐TC‐120 eight‐channel thermocouple input modules and one FP‐RTD‐122 resistance thermometer input module is used. **Table 3** shows the measurement uncertainties (within a 95% confidence interval) of the five‐hole probe meas‐ urements. These values contain an error due to the multi‐parameter approximation, random error and the systematic error of the pressure transducers. The difference between the positive and the negative direction is a result of the multi‐parameter approximation. The measurement uncertainties of the static pressure and the total pressure at test rig inlet as well as at stage inlet are ±1 mbar. Uncertainties for total pressure measurements up‐ and downstream of the TEGVs are also in the range of ±1 mbar. The overall uncertainty of the total pressure loss coefficient ζ is estimated to be about ±0.0014. The random fluctuation of rotational speed is below 0.2% of the current operating speed. Measurement uncertainty of the temperature measurement is about ±0.5 K. The day‐to‐day variation of the operating parameters such as pressure ratio, corrected speed, rotational speed, total pressure and temperature at rig inlet has been below 0.5%.


**Table 3.** Measurement uncertainties of the five-hole probe.
