Sea Level Static Condition Research Articles

Numerical Study on the Effect of Natural Gas on Aircraft Turbofan Engine Propulsion

The world has been concerned and worried about the depletion of the liquid oil fuel, besides new environmental rules are to be followed to reduce pollution hazards and global warming. The utilization of natural gas (as a near term fuel) and hydrogen (as long-term fuel) are receiving great attention, because they have less pollution effects. Since the aviation has a great deal in environmental pollution effects due to the cruise flight in the upper troposphere (supersonic aircraft) or in the lower stratosphere (subsonic aircraft) where most of the ozone concentrate, which helps in protecting the earth form ultra violet radiation. Therefore, the use of alternate fuel has a great attention in aviation. In the present study, the thrust specific fuel consumption and specific thrust for the aircraft during aircraft flight profile are predicted, when using aviation fuel and natural gas. The P&W JT9D –7R-turbofan jet engine is taken as a base line engine propelling the Boeing 747-200 aircraft as a base line aircraft with four engine nacelles mounted on wings. The model engine fuel-air cycle representation is carried out for design point calculations based on sea level static conditions and variable specific heats along engine components. The predicted engine performance results compared very well with the reported values by the manufacture. Predictions carried out using aviation fuel and natural gas show an increase in the specific thrust by 3% and decrease in the thrust specific fuel consumption by 14% and fuel to air ratio by 11%, when using natural gas.

Sea-Level Static Tests of Rocket–Ramjet Combined Cycle Engine Model

A rocket–ramjet combined cycle engine model, embedding twin rocket chambers driven with gaseous hydrogen and oxygen on the top wall side of a scramjet flowpath, was tested in its ejector-mode operation under sea-level static conditions. Gaseous hydrogen was also injected through secondary injector orifices to pressurize a ramjet combustor located downstream of the rocket chambers. Mixing between the hot rocket plume and cold airflow as well as combustion of residual fuel within the plume with the airflow caused entropy and static pressure increases in the constant-area mixing duct in the original flowpath design, resulting in a limited airflow rate. The mixing duct was redesigned to incorporate divergence from its onset to counter the pressure rise. With the diverging flowpath configuration, the airflow rate was increased by 40 percent. However, this modified flowpath geometry resulted in the formation of a low-speed area, through which the pressure rise in the ramjet combustor penetrated the mixing duct and reduced the incoming airflow rate. Without mechanical contraction near the engine exit, increasing pressure in the diverging portion of the flowpath was difficult, which in turn, resulted in poor mixing between the airflow and the rocket plume. Balancing these factors and additional mixing enhancement are required. The engine performance is then summarized.

Experimental study on the rocket-ejector system with a throat in the secondary stream

This study concentrated on hot-fire tests to demonstrate the flow characteristics that influenced by the throttle of the secondary stream in the rocket-ejector system. The experiment was conducted under the sea level static (SLS) condition. A small rocket used gaseous oxygen / liquid kerosene propellant as a primary stream supply system, and the high-pressure air bottles were employed as the secondary stream supply system. The equivalence ratio of the rocket, the air mass flow rate and the mixing length were varied to study the performance. The results show that there is a mass flow rate mismatch between the air flow entering through the inlet and the mass flow permitted due to aerodynamic choking. If the air mass flow rate through the throat is more, the inlet may unstart. If the air mass flow rate through the throat is less, the total pressure loss of the air due to the pseudo-shock would be caused under the effect of the suction performance. The unsuitable air mass flow rate is detrimental to RBCC engine operation in the ejector mode. With the increasing equivalence ratio of the rocket, the pressurization performance is improved accordingly, but the suction performance is obviously decreased. When the equivalence ratio is more than 1.1, the reacting shear layer would affect the suction performance.

Computational Fluid Dynamics Investigation of a Core-Mounted Target-Type Thrust Reverser—Part 1: Reverser Stowed Configuration

During the second half of the 90 s, NASA performed experimental investigations on six novel thrust reverser (TR) designs; core-mounted target-type thrust reverser (CMTTTR) design is one of them. To assess the CMTTTR efficiency and performance, NASA conducted several wind tunnel tests at sea level static (SLS) conditions. The results from these experiments are used in this paper series to validate the computational fluid dynamics (CFD) results. This paper is part one of the three-part series. Parts 1 and 2 discuss the CMTTTR in stowed and deployed configurations; all analyses in the first two papers are performed at SLS conditions. Part 3 discusses the CMTTTR in the forward flight condition. The key objectives of this paper are: first, to perform the three-dimensional (3D) CFD analysis of the reverser in stowed configuration; all analyses are performed at SLS condition. The second objective is to validate the acquired CFD results against the experimental data provided by NASA (Scott, C. A., 1995, “Static Performance of Six Innovative Thrust Reverser Concepts for Subsonic Transport Applications: Summary of the NASA Langley Innovative Thrust Reverser Test Program,” NASA—Langley Research Centre, Hampton, VA, Report No. TM-2000-210300). The third objective is to verify the fan and overall engine net thrust values acquired from the aforementioned CFD analyses against those derived based on one-dimensional (1D) engine performance simulations. The fourth and final objective is to examine and discuss the overall flow physics associated with the CMTTTR under stowed configuration. To support the successful implementation of the overall investigation, full-scale 3D computer aided design (CAD) models are created, representing a fully integrated GE-90 engine, B777 wing, and pylon configuration. Overall, a good agreement is found between the CFD and test results; the difference between the two was less than 5%.

Adaptation Method for Overall and Local Performances of Gas Turbine Engine Model

An adaptation method was proposed to improve the modeling accuracy of overall and local performances of gas turbine engine. The adaptation method was divided into two steps. First, the overall performance parameters such as engine thrust, thermal efficiency, and pressure ratio were adapted by calibrating compressor maps, and second, the local performance parameters such as temperature of component intersection and shaft speed were adjusted by additional adaptation factors. An optimization technique was used to find the correlation equation of adaptation factors for compressor performance maps. The multi-island genetic algorithm (MIGA) was employed in the present optimization. The correlations of local adaptation factors were generated based on the difference between the first adapted engine model and performance test data. The proposed adaptation method applied to a low-bypass ratio turbofan engine of 12,000 lb thrust. The gas turbine engine model was generated and validated based on the performance test data in the sea-level static condition. In flight condition at 20,000 ft and 0.9 Mach number, the result of adapted engine model showed improved prediction in engine thrust (overall performance parameter) by reducing the difference from 14.5 to 3.3%. Moreover, there was further improvement in the comparison of low-pressure turbine exit temperature (local performance parameter) as the difference is reduced from 3.2 to 0.4%.

On-design exergy analysis of a combustor for a turbojet engine: a parametric study

PurposeThe purpose of this study is to perform an exergy analysis of a turbojet engine combustor at different cycle parameters.Design/methodology/approachBase cycle parameters have been defined for the engine, and then differentiation of the combustor exergy efficiencies and destruction rates have been evaluated by changing overall pressure ratio, combustor exit temperature and combustor pressure ratio.FindingsFor the basic engine cycle, combustor unit is found to have lowest exergy efficiency as 62.3 for the sea level static condition. Compressor turbine exhaust and whole engine exergy efficiencies have been calculated as 88.7, 96.5, 68.2 and 69.4, respectively.Practical implicationsBecause of the biggest exergy, destruction is seen mainly in combustion system; effect of the combustor inlet pressure (related to the compressor design technology), pressure drop and exit temperature on the exergy efficiencies have been analyzed and combustor second law efficiency have been evaluated.Social implicationsThe investigation’s purposes are highly connected with social wellness and targeted at sustainable development of the society. Practical implementation of the obtained scientific results is directed on the improving of combustor for a turbojet engines and decreasing negative influence on the environment.Originality/valueAs a result of this paper, the following are the contribution of this paper in the field of gas turbine exergy subjects: Combustor has been found as the most critical component in respect of the exergy efficiency. Therefore, the effect of the combustor main cycle parameters such as inlet pressure, combustor pressure ratio and exit temperature have been analyzed.

Effect of Primary Rocket Jet on Thermodynamic Cycle of RBCC in Ejector Mode

AbstractThrust augmentation in ejector mode is important to improve engine performance so as to enable more propulsion applications of RBCC. Usually, the internal flow-path is configured mainly for combustion organization in ramjet and scramjet modes, thus the RBCC performance in ejector mode will mainly depend on the primary rocket operating parameters and jet expansion state. Considering such restrictions and requirements on the primary rocket, this paper studies the effects of the primary rocket jet on the thermodynamic cycle performance of ejector mode, in which incoming air is dominated by ejector-suction by the primary rocket jet at low flight Mach numbers. It is found that the engine performance in ejector mode will be improved by increasing the primary rocket chamber pressure and nozzle expansion ratio. Furthermore, for better design of primary rocket, primary rocket chamber pressure should be maximized on the basis of ensuring a complete expansion of primary rocket to the ambient pressure at the design point of flight conditions after 3-D CFD simulations for a full flow-path of RBCC including fore-airframe and aft-airframe of a flight vehicle. The results of 3-D numerical simulations show that the bypass ratio and specific impulse increase by 35.5 % and 12.5 % respectively at sea-level static condition.

Numerical Investigation of Shrouded Ejector-Enhanced Pulse Combustor Performance at High Pressure

A computational fluid dynamics investigation of a pressure-gain combustor system for gas turbine applications is described. The system consists of a valved pulse combustor and an ejector, housed within a shroud. The computational fluid dynamics solver is first validated by comparing its output to experimental measurements done on a pulse combustor operating at sea-level static conditions. Subsequently, the combustor inlet pressure and temperature are increased in the simulations to more closely align with those seen downstream of a compressor. Several parameters that influence combustor performance are then varied in order to demonstrate the potential benefit of optimization efforts. Emissions calculations are also performed. It is found that efficient high-pressure operation is possible and that system pressure gain of nearly 3% (at gas turbine compatible combustor temperature ratios) is achievable, with nearly smooth exit flow, having a competitive emission index. The potential for pulse combustors to achieve ultra-low levels of emissions at high-pressure conditions is also demonstrated.

Research on hypersonic aircraft using pre-cooled turbojet engines

Systems analysis of a Mach 5 class hypersonic aircraft is performed. The aircraft can fly across the Pacific Ocean in 2h. A multidisciplinary optimization program for aerodynamics, structure, propulsion, and trajectory is used in the analysis. The result of each element model is improved using higher accuracy analysis tools. The aerodynamic performance of the hypersonic aircraft is examined through hypersonic wind tunnel tests. A thermal management system based on the data of the wind tunnel tests is proposed. A pre-cooled turbojet engine is adopted as the propulsion system for the hypersonic aircraft. The engine can be operated continuously from take-off to Mach 5. This engine uses a pre-cooling cycle using cryogenic liquid hydrogen. The high temperature inlet air of hypersonic flight would be cooled by the same liquid hydrogen used as fuel. The engine is tested under sea level static conditions. The engine is installed on a flight test vehicle. Both liquid hydrogen fuel and gaseous hydrogen fuel are supplied to the engine from a tank and cylinders installed within the vehicle. The designed operation of major components of the engine is confirmed. A large amount of liquid hydrogen is supplied to the pre-cooler in order to make its performance sufficient for Mach 5 flight. Thus, fuel rich combustion is adopted at the afterburner. The experiments are carried out under the conditions that the engine is mounted upon an experimental airframe with both set up either horizontally or vertically. As a result, the operating procedure of the pre-cooled turbojet engine is demonstrated.

Reduction of NO emissions in a turbojet combustor by direct water/steam injection: Numerical and experimental assessment

Numerical and experimental investigations are conducted to assess the benefits and drawbacks of both water (mist) and steam direct injection within the combustion chamber of a 200 N static thrust turbojet. For this purpose, a three-dimensional CFD model of the combustion process is implemented where pollutant emissions are calculated; in parallel, a test campaign on the turbojet at sea level static conditions is carried out. In both cases the refrigerant flow is injected directly into the combustor, outside the liner. The aim of the investigations is to evaluate the impact of increasing water and steam flows (ranging from 0% to 200% of the fuel mass flow) onto the emissions levels (NO and CO) of the engine.

A feasibility study of a new ATREX engine system of aft-turbine configuration

A feasibility of ATREX (Air-Turbo-Ram Expander cycle) engine with conventional aft-turbine configuration has been studied to be developed in about 10 years, if the development project has started under enough resources. The novel tip-turbine of the original ATREX engine is replaced by a conventional aft-turbine, and the maximum turbine inlet temperature (TTT) is reduced to 1200K, to realize the engine by only using approved metal technologies of modern jet engines. The capability of the performance has been shown by parametric studies by changing components' design parameters. The study shows that the performance of the ATREX engine is not less than that of pre-cooled turbo jet. Some technical issues on developing the new ATREX engine have been addressed. The most important issue would come from the transient total temperature change due to the rapid acceleration from sea level static (SLS) condition (288K) to Mach 6 at 30km of altitude (1680K) in 6 minutes. The deformation due to transient thermal expansion has to be controlled. Especially, the change of the tip clearance and the clearance between rotors and stators are pointed out to be important design issues. The ATREX engine, which has shorter axial length and simpler rotor, has structural advantage over turbo jet.

Performance simulation of a turboprop engine for basic trainer

A performance simulation program for the turboprop engine (PT6A-62), which is the power plant of the first Korean indigenous basic trainer KT-1, was developed for performance prediction, development of an EHMS (Engine Health Monitoring System) and the flight simulator. Characteristics of components including compressors, turbines, power turbines and the constant speed propeller were required for the steady state and transient performance analysis with on and off design point analysis. In most cases, these were substituted for what scaled from similar engine components’ characteristics with the scaling law. The developed program was evaluated with the performance data provided by the engine manufacturer and with analysis results of GASTURB program, which is well known for the performance simulation of gas turbines. Performance parameters such as mass flow rate, compressor pressure ratio, fuel flow rate, specific fuel consumption and turbine inlet temperature were discussed to evaluate validity of the developed program at various cases. The first case was the sea level static standard condition and other cases were considered with various altitudes, flight velocities and part loads with the range between idle and 105% rotational speed of the gas generator. In the transient analysis, the Continuity of Mass Flow Method was utilized under the condition that mass stored between components is ignored and the flow compatibility is satisfied, and the Modified Euler Method was used for integration of the surplus torque. The transient performance analysis for various fuel schedules was performed. When the fuel step increase was considered, the overshoot of the turbine inlet temperature occurred. However, in case of ramp increase of the fuel longer than step increase of the fuel, the overshoot of the turbine inlet temperature was effectively reduced.

極超音速空気吸い込み式エンジン用予冷却器（プリクーラ）の開発研究

Here is presented an experimental and analytical study on a precooler for hypersonic air-breathing engines. Precooling of the incoming air breathed by an air-inlet gives extension of the flight envelope and improvement of the thrust and specific impulse. Three precooler models were installed into an air-turbo ramjet engine and tested under the sea level static condition. When the fan inlet temperature was down to 180K, the engine thrust and specific impulse increased by 2.0 and 1.2 times respectively. Thick frost formed on the tube surfaces at the entrance part of the precooler blocked the air-flow passage. On the other hand, very thin frost formed at the exit part because the water vapor included in the air was changed to mist particles due to the low temperature of the air in this part. Parametric studies on the precooler design values and a sizing analysis were also performed. Decrease of tube outer diameters on the precooler is only way to increase heat exchange rates without increase of its weight and pressure loss.

フレームホルダ二段化によるラム燃焼器の小型化に関する研究

In order to carry heavier payloads and to fly longer distance, the compact engine is required. The ram combustor with the double-staged flameholders was introduced to shorten the ram combustor which is a major component of the ATR. An additional flamefolder is placed downstream of the conventional (single-staged) one to achieve quick flame expansion in the downstream direction. The ram combustion test was carried out in a 120 deg sector chamber for three cases of flameholder configuration to investigate the various effects on the combustion and aerodynamic performances. The ram combustor with the double-staged flameholders increased the combustion efficiency without large total pressure loss and therefore it was able to shorten 40% axial length of the ram combustor in comparison with the ram combustor with a conventional flameholder. In this study, for the compact ram combustor, the combustion efficiency is greater than 80% on sea level static condition where the air temperature and pressure at the combustor inlet is much lower than that of the supersonic flight condition.

小型・高推力エアーターボラムエンジン

The Air Turbo Ramjet (ATR) is a combined cycle engine which performs like a turbojet engine at subsonic speeds and a ramjet at supersonic speeds and therefore the ATR is an attractive propulsion system for the wide operation range (e.g. Mach 0 to Mach 4). The ATR can provide a higher specific impulse than a solid fuel rocket engine and a higher thrust per frontal area than a turbojet engine. The major ATR components are the inlet, fan (compressor), turbine, gas generator, combustor and exhaust nozzle. In the ATR, the turbine drive gas is generated by a decomposed liquid or solid fuel gas generator. In order to carry heavier payloads and to attain shorter flight time, the compact and high thrust engine is required. In this study, the ram combustor with the double-staged flameholders and the fan with tandem blade were introduced to shorten the engine length and to increase the fan pressure ratio, respectively. Furthermore, the engine testing was carried out on sea level static condition to confirm the engine component integration technologies for the ATR propulsion system.

Development Study of a Precooler for the Air-Turboramjet Expander-Cycle Engine

A review of the development of a precooler for the air-turboramjet expander-cycle (ATREX) engine is given. Three types of precooler for the ATREX engine ground-testmodel were designed, manufactured, and tested under sea-level static conditions. The results suggested two problems affecting the precooler performance, heat transfer rate and air ow pressure drop. One is nonuniformity of the air ow through the tube banks. The other problem is frost formation on the heat transfer surfaces. Concerning nonuniformity of air ow, the shell conŽ guration was modiŽ ed based on analysis by computational  uid dynamics calculation. To improve the precooler performance under frosting condition, a new method to add a condensable gas into the air ow was proposed and examined by experiments on a subscale heat exchanger model. Addition of a small quantity of ethanol can effectively restrain the decline of the precooler performance due to frost formation.

Model-based reasoning for fault diagnosis of twin-spool turbofans

A simplified rule-based expert system can be of great use to maintenance engineers involved in on-line health monitoring of aero-engines. The development of a rule base has been done on the following lines: 1 A design and off-design mathematical model of the turbofan has been developed. 2 Healthy and fault-implanted engines have been run at various off-design conditions to see the effect of faults on certain aero-thermodynamic performance parameters, with appropriate selection of independent parameter(s). 3 From these calculated values, fault matrices have been developed for sea-level static conditions taking the net thrust as an independent parameter. 4 From these fault matrices, rules have been developed which form the knowledge core of the expert shell. These rules have been developed for Garrett TFE 731-2, a moderate bypass and overall pressure ratio, generic twin-spool turbofan and so the package (off-design code plus expert system) can serve as a pedagogical tool for training of engineers in the aero-engine industry and academic institutions.

Development Study on ATREX Engine

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Institute of Space and Astronautical Science (ISAS) cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the combined cycle (a fan-boosted ramjet) engine providing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling the incoming air as shown in Fig. 1 . ATREX employs the tip turbine configuration which allows the compactness and the light weight of turbo machinery and the variable geometry airintake and plugnozzle which allow the wide range operation conditions.From 1990 to 1992, “ ATREX-500“ has been tested at the sea level static conditions. ATREX-500 is the 1/4-scale model of which fan inlet diameter is 300mm and overall length 2,200mm. From 1992 have been performed the wind tunnel tests on the primary components of ATREX, the axisymmetric variable geometry airintakes, the precoolers and the variable geometry plug nozzles. In parallel to the windtunnel tests, the ram combusters have been tested simulating the hypersonic flight conditions and the application studies on advanced carbon-carbon composite for the tip-turbine and fan assembly has been proceeded.In 1994 initiated the flight test plan in which ATREX will be verified in the practical flight conditions by using an unmanned flying test bench.In 1995 will be tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing on the precooler.The present paper addresses the high loading ram combuster experiment using the mixer with skewed lobes to generate swirl flow and the analytical studies and the designs on the precooler and the precooled ATREX engine and the flight test plan.

Application of h ∞ robust control to the RM12 jet engine

Turbojet engines are becoming increasingly complex, with more control variables, to meet future demands on performance, fuel consumption and functionality. To utilise the potential of the engines, it is also necessary to use more advanced control concepts than are conventionally used today. The trend is towards control concepts often referred to as ‘Smart Engines’. Multivariable controllers are most often the core of these systems. A multivariable control concept for a jet engine is proposed and investigated in this paper. The feedback control law is based on H ∞-design. The control concept is applied to the RM12 jet engine for the Swedish fighter aircraft JAS 39 Gripen in a research program for future advanced control. The controller has so far been developed for the core engine operating range, at sea-level static conditions, and evaluated via simulation with a detailed non-linear model. This evaluation clearly shows the potential of using more advanced control methods in the area of jet engine control. Finally, suggestions on how to continue the research program towards a future advanced control concept are given.

Development study on ATREX engine

This is the status report of the development study on ATREX engine (Air Turbo Ramjet) that is now under way in the Instutute of Space and Astronautical Science (ISAS) in cooperation with the Ishikawajima Harima Heavy Industries (IHI), the Kawasaki Heavy Industries (KHI), the Mitsubishi Heavy Industries (MHI). ATREX engine will be applied for the propulsion system of fly-back booster of TSTO space plane. ATREX is the fan-boosted ramjet producing the effective thrust from sea level static to flight Mach number 6. ATREX is worked on the expander cycle with precooling incoming air as shown in Fig.1. ATREX employs the tip turbine configuration which allows compactness and light weight of turbo machinery and the variable geometry airintake and plug nozzle which allows the wide range flight conditions. ATREX development study has been conducted with the sea level static tests since 1990. ATREX engine for tests is the scaled model designated by “ATREX-500” of which fan inlet diameter is 300 mm and overall length 2,200 mm. From 1992 have been performed the wind tunnel tests on the primary aerodynamic components such as the axisymmetric variable geometry air intakes, the precoolers and the variable geometry plug nozzles. The application study on advanced carbon-carbon composite for the tip-turbine and fan assembly has been conducted. This study status is presented in the another session of this IAF congress 1. The flight test of ATREX is now planning to verify the engine performance and functions in the practical flight conditions by using an unmanned flying test bench. In 1995 was tested ATREX-500 installing the precooler under the sea level static conditions to examine the engine performance and the icing problem on the precooler. The present paper addresses primarily the test results of precooled ATREX engine.