Pressure And Temperature Distortion Testing Of A Two

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THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 345 E. 47th St., New York, N.Y. 10017 93-GT-118 The Society shall not be responsible for statements or opinions advanced in papers or discussion at meetings of the Society or of its Divisions or Sections, or printed in its publications. Discussion is printed only if the paper is published in an ASME Journal. Papers are available from ASME for 15 months after the meeting. Printed in U.S.A. Copyright © 1993 by ASME PRESSURE AND TEMPERATURE DISTORTION TESTING OF A TWO-STAGE CENTRIFUGAL COMPRESSOR W. T. Cousins, K. K. Dalton, and T. T. Andersen Performance and Operability AlliedSignal Propulsion Engines Phoenix, Arizona G. A. Bobula U.S. Army Vehicle Propulsion Directorate Army Research Laboratory Cleveland, Ohio ABSTRACT Altitude pressure and temperature inlet distortion testing of the two-stage centrifugal compressor in the T800-LHT-800 engine is described. The test setup and the testing techniques are reviewed and the results of the test are presented. The generation of classical 180 degree patterns of both pressure and temperature distortion is discussed. Temperature distortion was created using a hydrogen burner system while pressure distortion was created in the classical manner, using screens. Results of both individual and combined temperature and pressure distortions in both opposed and concurrent patterns are shown. NOMENCLATURE N PAV PAVLOW TAV TAVHI We ,613 c/P ATc/T 5 0 physical speed average pressure average pressure in the low-pressure sector average temperature average temperature in the high-temperature sector corrected flow, W'0/8 circumferential pressure distortion intensity circumferential temperature distortion intensity normalized pressure normalized temperature INTRODUCTION Characterizing the effects of inlet airflow distortion on the stability of an aircraft engine is a critical part of any aircraft engine development and qualification program. Understanding the response of the engine compression system to these destabilizing influences enables proper matching of the engine with the aircraft. Throughout an engine program, many influences on engine stability are evaluated. These are used in the formulation of an engine stability assessment. The stability assessment contains all the destabilizing influences of the engine, along with mission analysis information for the particular aircraft. Inlet distortion is a major part of the stability assessment. The engine response to inlet distortion is typically obtained through extensive testing, both at sea level and altitude conditions. Inlet flow distortion can be separated into two types: pressure and temperature. Often, pressure distortion is considered while temperature distortion is not. This is usually because temperature distortion is more difficult to simulate under test conditions, and depending upon the aircraft mission (for example, civil aircraft), the probability of encountering temperature distortion may be very small. This does not mean that temperature distortion is not important. Much to the contrary, temperature distortion can often have a greater effect on engine stability than pressure distortion. In military applications, the effect of temperature distortion is usually quite critical. Temperature distortion is typically generated by weapons firing, steam ingestion on aircraft carriers, and the exhaust from other aircraft. In rotorcraft, it may even result from ingestion of one's own exhaust during flair and hover. Generating pressure distortion with screens in a test cell is usually easier than generating temperature distortion. Consequently, pressure distortion testing is a common practice while temperature distortion testing is performed less often. Pressure distortion is due to nonuniform inlets, aircraft maneuvers, crosswinds, inlet/fuselage boundary layers, etc. Often in flight, pressure and temperature distortion are present simultaneously. The effect of this combined distortion can be different from the effect of each separately, therefore, tests were conducted to examine the combined phenomena. Presented at the International Gas Turbine and Aeroengine Congress and Exposition Cincinnati, Ohio May 24-27, 1993 This paper has been accepted for publication in the Transactions of the ASME Discussion of it will be accepted at ASME Headquarters until September 30,1993 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use TEST FACILITY This paper summarizes the altitude inlet distortion testing that was performed on the T800-LHT-800 (T800) engine to examine the effects of pressure and temperature distortion on the compression system surge margin at various altitudes. Testing was conducted to acquire data for the development and qualification of the T800 engine, during the time period from April to August 1991. This data is used in the formation of a stability assessment and performance computer model, thus providing the ability to help customers evaluate the effect of inlet airflow distortion on both performance and engine compression system stability around the aircraft flight envelope. The inlet distortion test was conducted in test cell number 3 of the Propulsion Systems Laboratory (PSL3) at the NASA Lewis Research Center. PSL3 is an air breathing engine altitude facility capable of simulating the environment for full-scale engines at altitudes up to 70,000 feet (21,336 m) and at velocities up to Mach 3. The altitude tank is 24 feet (7.32 m) in diameter and 38 feet (11.58 m) long (Figure 2). Temperature and pressure conditioned air is supplied to the inlet of the altitude chamber from the central air delivery system. A large bulkhead isolates the air delivery side of the test cell from the altitude chamber. Normally, the test engine inlet is connected This work was performed at the NASA Lewis Research Center in Cleveland, Ohio, with the support of the Army Aviation Systems Command. The T800 is being developed for the U.S. Army by the Light Helicopter Turbine Engine Company (LHTEC), a partnership between Allison Gas Turbine Division of General Motors and AlliedSignal Propulsion Engines. TEST VEHICLE The T800 (Figure 1) contains a two-stage centrifugal compression system coupled to a two-stage gas generator turbine with a two-stage freeshaft power turbine. The engine inlet is annular in nature and incorporates an integral inlet particle separator (IPS). The T800 is of the 300-1b, 1200-shaft horsepower class. For this test series, the full engine was tested with the power turbine coupled to a dynamometer system. An inflow-bleed system was used to set the pressure ratio of the compression system to obtain data off the normal operating line. Data was obtained along constant speed lines from the operating line to surge. Fig. 2 NASA Lewis Research Center Propulsion System Laboratory Test Cell #3 Fig. 1 The T800-LHT-800 Engine directly to this bulkhead, with a bellmouth on the upstream side, and the engine, downstream of the bulkhead, subjected to simulated altitude conditions. Altitude conditions are maintained in the chamber by controlling the flow from the test cell to the central air services building. For the T800 installation, the engine could not be mounted directly on the bulkhead since the power turbine shaft exits at the front of the engine and needed to be connected to the dynamometer. To accommodate this configuration, the test cell bulkhead was blanked off and a pipe with a small bellmouth inlet was connected to the bulkhead plate. This lead to flow control valving, flow measuring hardware, an inlet settling chamber or airbox, and the engine as shown in Figure 2. The engine and airbox were mounted directly on a large pallet. This pallet provided the rigid base of a removable test installation upon which the major experimental hardware, including the engine, airbox, and dynamometer were mounted. Using this palletized stand, it was possible to assemble and work on significant portions of the test vehicle before occupying the test cell. 2 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use pressure and temperature probes (ten rakes with four probes each equally spaced on centers of equal area) at the engine aerodynamic interface plane (AIP). Ten static pressure taps were also arranged on the shroud in this location. SPECIAL TEST HARDWARE The main items of special test hardware were those related to imposing the inlet flow distortions on the engine. Pressure distortions were generated by inserting a 180 degree circumferential extent screen in the flowpath upstream of the inlet. To raise the velocity at the screen (and therefore the pressure drop through it) the screen was located in a convergent-divergent inlet duct section. The screen was composed of an overlay of two separate screens and was mounted on a large mesh structural grid, or backer screen (Figure 3). These were located in a rotatable spool piece that permitted circumferential clocking of the screens to enable detailed definition of the resulting distortion pattern with a fixed array of inlet rakes. Compressor impeller inlet instrumentation consisted of five rakes with seven total pressure elements per rake, and five shroud static taps. Other static taps were located at the front frame flowpath exit. The compressor exit was instrumented with eleven rakes (each having three total pressure and two total temperature elements), in addition to ten plenum static pressure taps. Fig. 4 Hydrogen Burner Used to Generate Temperature Distortion Dynamic data was acquired through high-response pressure transducers (Ku lite s) and high-response tern peratu re thermocouples. This instrumentation was located at the compressor inlet, between the stages, and the compressor exit. The dynamic data was used to assure that the distortions being generated did not produce any separation regions in the inlet particle separator and to examine details of the surge behavior if necessary. Since no anomalies were found in the engine, this data is not discussed in this paper. Fig. 3 180 Degree Distortion Screen Mounted on the Backer Screen The temperature distortions were generated using a gaseous hydrogen burner (Figure 4) located in the duct ahead of the pressure distortion screen holder. The burner design was based on the work of earlier NASA/Army Propulsion Directorate inlet distortion research (Klann, et al., 1984). For this test, modifications based on prior T800 preliminary flight release inlet distortion testing and subsequent government improvements aimed at developing this test device for small engines with annular inlets were incorporated. The burner was divided into six sixty-degree segments, allowing the generation of distortions in multiples of sixty degrees, including multiple-per-rev patterns. The T800 distortion testing reported herein used only 180 degree circumferential distortion patterns. TEST OBJECTIVES The overall objective of the T800 altitude distortion test was to examine compression system stability and gather data to better understand the T800 operability characteristics under both pressure and temperature distortion. The specific objective was to document the changes in surge margin under distorted inflow conditions, at three altitude conditions, so the sensitivity models could be built and used in the stability assessment for various installations. These conditions were sea-level static, three kilometers (9,843 ft.) Mach 0.6, and nine kilometers (29,528 ft.) Mach 0.2. INSTRUMENTATION The T800 engine was a fully instrumented test vehicle. The inlet airflow distortion level was defined using an array of 40 total 3 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use on the characteristics of the compressor. The effect of the temperature and pressure distortion is then compared to the backer map. In later analyses, the changes due to distortion are analytically applied to the "clean" configuration (in other words, the effects of the hardware on the compressor are removed). For purposes of this paper, the data is shown with respect to the backer map, as it was taken. Four distortion configurations were tested: pressure distortion alone, temperature distortion alone, a concurrent pattern of pressure and temperature distortion, and an opposed pattern of pressure and temperature distortion. All of the patterns tested in this test series were 180 degree distortions. This was to provide the compressor with the most adverse distortion possible. Using 180 degree distortion patterns allows a passing impeller blade a maximum residence time in the low momentum region of flow. This is the worst case distortion for a machine of this type, since the aerodynamic time constant associated with blade stalling and recovery is a function of blade chord. Incorporating centrifugal impellers, the time constant of the T800 compression system is significantly longer than that of a compression system with axial blades. This is why centrifugal compressors are typically more tolerant to distortion than axial compressors. In addition, dynamic distortion was not investigated with this centrifugal impeller system, due to the aerodynamic time constant. 3 km, Mach 0.6 20 18 — 16 — 15 14 - Temperature changes over time (temperature ramps), while important to the overall operability of the engine, were not run in this test series, since tests of this nature were performed earlier in the engine development program at the NASA facility with no adverse results, while running to the engine specification limits. 64— 2 1 0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 Compressor Corrected Flow, WITA5 , Ibm/sec RESULTS Baseline Determination Fig. 6 Clean 3km., Mach 0.6 Map To determine the baseline operating characteristics of the compressor, data was obtained to develop the "clean" map. Figures 5, 6, 7, and 8 show the clean maps at the three altitude conditions tested, along with the available constant speed surge margin. Next, all of the distortion hardware used to hold the screens and burners was installed and the "backer" map developed (Figure 9). This was necessary since the hardware has an effect 9 km, Mach 0.2 NASA Ground Level 1 0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 Compressor Corrected Flow, Wil3715 , Ibm/sec Fig. 7 Clean 9km, Mach 0.2 Map 3.0 4.0 5.0 6.0 7.0 Prior to the end of the test series, a rub occurred on the second stage impeller. Believed to be a result of running at a high inlet Mach number condition (3km, Mach 0.6) with temperature distortion, it had the effect of reducing the efficiency on the second stage. A second backer map' was obtained, so that subsequent distortion data could be properly baselined. The data 8.0 Compressor Corrected Flow, WK'S, lbm/sec Fig. 5 Clean NASA Level Map 4 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use (PAV) (PAVLOW) Circumferential Location, degrees APe (PAV) - (PAVLOW) P (PAV) 0 90 70 80 60 Compressor Corrected Speed (%) 50 100 Fig. 8 T800 Compression System Maintains Acceptable Available Surge Margin Over the Flight Envelope (TAVHI) (TAV) NASA Ground Level — - Backer Stability Limit & Op-Line Prior to Rub 18- • Backer Map Data Prior to Rub — Clean Map Circumferential Location, degrees 20 (TAVHI) (TAV) ATc T (TAV) Fig. 10 Distortion Intensity Definition Used in the Test Series • NASA Level • NASA 3 km. Mach 0.6 • NASA 9 km. Mach 0.2 2.00 47 2 1.50 60 10 2.0 3.0 4.0 5.0 6.0 7.0 ■ t I e • AA 8.0 -o ° Compressor Corrected Flow, WIFA5, Ibm/sec Fig. 9 Effect of Distortion Hardware on the Compressor Characteristic 0 1.00 -4.0 • • (Opposed) 0.50 0.00 presented in this paper is always compared with the appropriate backer map. • -2.0 2.0 0 0 ATc/T 4.0 60 )2 0.05 ( Wc/ Wcdesign The distortion testing procedures/analysis techniques performed for the T800 engine follow the practices of the Society of Automotive Engineers (SAE) S-16 Committee on Inlet Distortion (1978, 1983, 1989). These documents describe the convention used in this test series to describe "distortion intensity," APc/P and ATc/T, as shown in figure 10. Pressure Distortion - NASA Ground Level As previously mentioned, the test covered temperature and pressure distortion alone and combined, in both concurrent and opposed configurations. Figure 11 shows the wide range of testing that was performed. The effect of pressure distortion on the engine was tested using a 180-degree, 1/rev circumferential wire mesh screen. Previous tests had shown that it was difficult to obtain the desired levels of total pressure loss due to a low inlet Mach number. To increase the pressure loss caused by the screen, the screen was positioned Fig. 11 Tested Distortion Intensities Exceeded the Engine Specification Limits 5 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use Temperature Distortion - NASA Ground Level in a converging-diverging section, just upstream of the AIP. The distortion screen produced a distortion intensity (APc/P) of 4.1 percent at a corrected flow of 8.5 lbm/sec (3.86 kg/sec). Before the temperature distortion testing at NASA Lewis Research Center, testing had been performed in Phoenix. The Phoenix testing used a hot air injection system to create the distortion, as reported by Cousins, et al. (1991). The hot air injection system was not as flexible as the NASA hydrogen burner, so the Phoenix test was run with a constant AT across the inlet of 75 degrees Fahrenheit (23.9 degrees Celsius) while the NASA test was run at a constant flow-normalized distortion intensity. Figure 13 shows a comparison of the resulting change in surge margin (due to the surge line movement only) for the two tests. Figure 14 shows the compressor characteristics from the NASA test when under the influence of temperature distortion. At the NASA ground level condition (the closest to sea-level static conditions that could be obtained in the test cell, 14.073 psia [97.03 kPa] Mach 0.08), the distortion caused the compression system surge line to shift to a lower flow, as shown in figure 12, below speeds of 92.5 percent. 20 18 NASA Ground Level ---- Pressure Distortion Stability Limit & Op-Line • Pressure Distortion Data — Backer Map Prior to Rub 16 105 103 100 97.5 14 Stability Limit 12 10 - Aar 95 92.5 o 6- 0.05 ( Wc /Wc design )2 • C Cs 0 V0 c ED a) 0 • -5 • 65 • 2 10 2.0 3.0 4.0 5.0 Compressor Corrected Flow, 6.0 7.0 KO, — 1.0 c E 5 al a) .c > ..4110'" 7o „Aro- 4- Constant 15 c 0 10 N/10 AMP 90 85 40r_80 1 4111.0"75 — Operating Line 8 NASA Ground Level Arc / T 20 8.0 m Ibm/sec te -10 ‘5 y -15 Constant AT = 75°F • NASA LeRC • Phoenix -20 50 Fig. 12 Pressure Distortion Has No Adverse Effect on 60 70 80 90 100 Corrected Compressor Speed (%) Compressor Stability Margin in Critical Low-Speed Areas Fig. 13 Different Surge Margin Results Between Phoenix and NASA Tests are a Result of Different Temperature Conditions In addition, the operating line also shifted to a lower flow but the magnitude was smaller than the surge-line shift, resulting in a net gain in surge margin at low speed. At speeds above about 92.5 percent, pressure distortion results in a loss in surge margin. This loss in surge margin at high speed is not critical, since the engine operational transients at these speeds do not use as much of the available surge margin as at lower speeds. The operating line did not change with temperature distortion, although the intersection of the speed lines with the operating line did change at high speed. Running each of these tests in a different manner allowed covering a wide range of inlet temperature levels that are typical of gun gas ingestion (figures 15 and 16). Both tests, however, showed that the effect of temperature distortion on the compressor is greater than the effect of pressure distortion. Prior to this test, distortion testing was performed at the AlliedSignal Propulsion Engines facility in Phoenix, Arizona (Cousins, et al., 1991), using the same distortion screens. Similar results were obtained in the Phoenix test. Opposed Temperature and Pressure Distortion - NASA Ground Level With the pressure distortion screens in place, temperature distortion was applied to the opposite side of the inlet from the pressure distortion. As shown in figure 17, both the stability limit line and the operating lines moved a significant amount. Only at 92.5 percent speed was there a surge margin decrease. As expected, distortions run in this manner do not show as adverse an influence on the compression system as either pressure or temperature by themselves. The stabilizing effect of the circumferential pressure distortion is not unlike that seen on some axial-flow machines, where circumferential and radial distortion affect the compression system in opposite ways. While this phenomenon is not well understood, it is possible that the forced, repetitive change in the momentum of the entering flow has a stabilizing effect on the inducer of the centrifugal impeller. 6 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use • 4.0 NASA Ground Level - - Temperature Distortion Stab. Limit & Op-Line 18 - • Temperature Distortion Data - Backer Map Prior to Rub 16 - 20 cNi g, 3.0 15 et, r 14) E 12▪ 10 8- O 6- 1 0 2.0 4.0 3.0 5.0 6.0 7.0 8.0 Fig. 14 Temperature Distortion Reduces the Compression System Stability Limit 1.0 • . 80 op 90 srg • 90 3 0 4.0 6.0 7.0 5.0 Corrected AIP Flow (Ibm/sec) • op 8.0 NASA Ground Level - Opposed Distortion Stab. Limit & Op-Line 18 - • Opposed Distortion Data - Backer Map After the Rub 20 NASA Ground Level 100 16 Z 14- 90 'E _c . 80 E 12 ♦ 70 'E 60 10 Typical Range of Gun Gas ingestion r 50 < a: 4- • 30 86- 40 60 2 10 20 10 0 .70 op • 80 srg Fig. 16 Flow-Normalized Temperature Distortion Intensity is Different Between Phoenix and NASA Tests Compressor Corrected Flow, WO, Ibm/sec 0 2.0 0.0 42 • NASA LeRC • Phoenix srg: surge point op: operating line point • 70 srg 60 70 80 90 3.0 4.0 5.0 , • , 6.0 7.0 8.0 Compressor Corrected Flow, Wir3/5 , Ibm/sec • NASA LeRC • Phoenix 50 2.0 Fig. 17 The Effects of Opposed Temperature and Pressure Distortion Are Not as Great as Temperature Distortion Alone 100 Compressor Corrected Speed (%) Fig. 15 Temperature Distortion Tests Cover a Range of Operating Conditions conditions, NASA ground level, 3 km Mach 0.6, and 9 km Mach 0.2. Figure 19 shows the surge margin calculated from the previously presented maps. Concurrent Tem peratu re and Pressure Distortion - NASA Ground Level For the NASA level testing, pressure distortion begins to adversely affect the stability of the compressor above about 90 percent speed, with a surge margin decrease of about 6 percent. The most critical region for operation of the engine is in the 70 to 85 percent speed, where transient accelerations require the most surge margin. In this region, pressure distortion enhanced stability, showing the compressor to be very tolerant of pressure distortion. This is similar at the other altitude conditions, Applying the temperature and pressure distortion in the same 180 degree sector results in an effect on the compression system that is a combination of each separately. The temperature distortion degrades stability while the pressure distortion enhances it, resulting in a net change shown in figure 18. Surge Margin Determinations - All Altitude Conditions Figures 19, 20, and 21 show the surge margin changes that were measured for all of the distortion patterns at the three altitude although at 3 km Mach 0.6 there was also a low speed loss in surge margin of about 5 percent. Examination of the compressor 7 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use • NASA Ground Level --- Concurrent Distortion Stab. Limit & Op-Line 18 - • Concurrent Distortion Data — Backer Map After the Rub 16 - n c 12 Stability Limit 10• 8 6 15 • 10 C 103 100 97.5 14 - 90 95 92.5 .c 5 • 2 Nne 80 - 75 4 es 2.0 0 co Operating Line - 20 4.0 5.0 6.0 7.0 50 8.0 Fig. 20 Distortion Causes a Shift in Both the Operating Line and the Stability Limit Fig. 18 Concurrent NASA Ground Level co 15 • 0 10 C 5 co 2 0 co .5 Cl) -10 4 4 • A • • Pressure -15 Distortion x Opposed Distortion 60 ; • • ij Y • • • Pressure Distortion • Temperature Distortion • Concurrent Distortion 60 70 80 90 100 Temperature and Pressure Distortion Have Similar Effects Above 80% Speed at 3 km, Mach 0.6 Conditions 9km, Mach 0.2 • ♦ 70 • ▪ 10 0 5 2 X • * • a • • Temperature Distortion 90 0 A• • -5 • -10 -15 • Concurrent Distortion 80 Cco C -20 50 • 15 to 20 C • x Opposed Distortion 20 • 25 • X Corrected Compressor Speed (%) Compressor Corrected Flow, Wri")/8, Ibm/sec 30 • 10 -15 70 3.0 )< X eC 2 10 • a _5 85 - 3km, Mach 0.6 20 20 • Pressure Distortion • Temperature Distortion x Opposed Distortion • Concurrent Distortion -20 100 50 Corrected Compressor Speed (%) 60 70 80 90 100 Corrected Compressor Speed (%) Fig. 19 Temperature Distortion Causes the Most Adverse Effect on Surge Margin at Low Altitude Fig. 21 Concurrent Distortion Generates the Largest Surge Margin Degradation at High Altitude map for this condition (Figure 6) shows that the stability limit line drops at low speeds, causing the low speed surge margin effect with pressure distortion. This is not much concern because if one were to operate at a 3 km Mach 0.6 condition, the engine would have to be at high power and not operating in this portion of the map. Concurrent distortion had the greatest adverse effect at 90 to 95 percent speed, using as much as 9 percent surge margin at the 9 km condition, but where available surge margin is 25 to 30 percent. Surge margin increases were obtained at other speeds. Similar results were obtained for the opposed distortion patterns. Generally, the opposed distortion patterns did not affect the surge margin to the magnitude of the concurrent distortion patterns. Temperature distortion has the most consistent and adverse effect on the compression system, losing about 3 percent surge margin in the critical map area around 70 to 85 percent speed. From the previous Phoenix test, this was also the case but the magnitude of the surge margin degradation was larger (approximately 5 percent), due to the higher temperature distortion intensity levels run at low speeds. In general, pressure distortion causes an increase in the stability characteristics of the engine at low speed and a slight decrease at high speed, while temperature distortion seems to cause a decrease overall. Comparisons of the distortion results presented in this paper with the results of the Phoenix test series (Cousins, et al., 1991) suggest that the overall effect of the combined patterns is 8 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use a function of both the level of each type of distortion and the particular pattern position (concurrent or opposed). SAE S-16 Technical Committee, 1991, "A Current A ssessment of the Inlet/Engine Temperature Distortion Problem," ARD50015. SUMMARY The inherent stability of a two-stage centrifugal compression system has been demonstrated. Testing has shown the T800 engine to be very tolerant of inlet airflow distortion, losing the most stability where there is the most available surge margin. This provides the engine with the capability to be installed in many different rotorcraft applications. The effect of pressure distortion on compression system stability is minimal and in most cases, causes an increase in stability margin. Temperature distortion reduces the stability of the engine, but the largest degradation is in regions of the compressor map where the most stability margin exists. Patterns of combined pressure and temperature distortion affect the stability in different ways, depending upon the screen orientation and the altitude conditions. Although this is the case, the combined distortion patterns had a minimal adverse effect on engine stability. ACKNOWLEDGEMENTS The authors wish to thank Brian Takesuye, Christopher Gromek, Tor Henriksen, Joan Doherty, Merritt Thomas, and Donald Palmer of AlliedSignal Propulsion Engines for their extended efforts and support of this test series. Also, thanks to David Clark and Teresa Kline of the Army Vehicle Propulsion Directorate; and Martin Ginley and Jeffrey Balser of the NASA Lewis Research Center for their dedicated contributions to the performance of this test. It is the effort of these individuals that has made this work possible. The authors would also like to thank LHTEC for permission to publish these test results, and the NASA Lewis Research Center and the Army Vehicle Propulsion Directorate for the support of this test series. REFERENCES Cousins, W.T., Miller, R.E., and Dalton, K.K., 1991, 'Distortion Tolerance of the T800-LHT-800 Turbosheft Engine," Proceedings, American Helicopter Society 47th Annual Forum and Technology Display, Vol. 2 (A92-14326 03-01) pp. 1147-1155. Klann, GA., Barth, R.L., and Biesiadny, T.J., 1984, 'Temperature Distortion Generator for Turboshcft Engine Testing," NASA TM-83478. SAE S-16 Technical Committee, 1978, "Gas Turbine Engine Inlet Flow Distortion Guidelines," ARP 1420. SAE S-16 Technical Committee, 1983, "Inlet Total-Pressure Distortion Considerations for Gas Turbine Engines," AIR1419. 9 Downloaded From: https://proceedings.asmedigitalcollection.asme.org/ on 09/13/2017 Terms of Use: http://www.asme.org/about-asme/terms-of-use