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EAS 201 Aerospace Propulsion Assignment Sample SUSS SingaporeEAS 201 Aerospace Propulsion Assignment Sample SUSS Singapore

EAS 201 Aerospace Propulsion Assignment Sample of SUSS, Singapore

In this assignment sample, we will be going to discuss the Aerospace Propulsion SUSS, Singapore. The aim of this course is to equip students with adequate knowledge about aerospace propulsion, a required skill for those looking into the aviation industry.

The majority of modern aircraft are powered by gas turbine engines and so this class introduces you to various methods before focusing on them more in-depth. You’ll learn how engineering principles apply when it comes to calculating engine thrusts as well as other necessary parameters that come up during theory work.

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The coursework for this subject will cover a wide range of topics from the construction and design of gas turbines to how they function in real-world conditions. There will be laboratory sessions that simulate different operating parameters including fuel and ignition systems, where students can learn more about their behavior under these settings while exploring computer simulations.

This course was designed to give the student a full understanding of the operation and design of gas turbine engines within the context of electrical power plants. “We teach students how to design, operate, maintain, and repair the equipment that is used in these gas turbines,” Dr. Okoro explained.

These systems are currently used in most electricity-producing power plants. They cover a range of sizes which allows each student to gain real-world experience as they progress through their coursework. It’s not just about theory without any hands-on work; Purdue currently has nine labs available for students enrolled in this program to get some real-life skill sets involving these systems.

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TOA, TMA, GBA Assignment solution of  EAS 201 Aerospace Propulsion modules SUSS, Singapore

At the end of this course, Singaporean students will be able to learn an Aerospace Propulsion module with the help of the following learning outcomes :

1.Explain the theory of gas turbines and ramjets.

Gas turbines are rotary engines that generate mechanical power from a working fluid that has been compressed to high pressure in an airtight combustion chamber and ignited. The rapid expansion of the hot gases produced by ignition causes the rate of rotation to increase in proportion to their temperature. Gas turbine engines produce more power for their weight than any other known type of internal combustion engine.

Therefore, gas turbines are used wherever there is a demand for relatively large amounts of mechanically generated power, particularly in aerospace propulsion systems. They are also widely used for electricity generation where efficient conversion of heat into useful work is required (e.g., industrial gas turbines).

Ramjet theory is a term used to describe the process by which a vehicle powered by a ramjet engine flies through the air. Ramjets work on the principle of airflow being injected into an inlet at high speed, mixed with fuel, and combusted to produce thrusts.

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The velocity of the airflow determines how fast combustion takes place inside the chamber (i.e. burn rates), and consequently governs exhaust gas velocities also (i.e. what frequency it will release energy) so that when one considers transmitting fusion reactions.

Ramjets work by diffusing air or some other reactive gas into the combustion chamber and mixing it with fuel, usually a hydrocarbon jet fuel, a kerosene derivative. The resultant exothermic synthesis produces heat and converts chemical energy to kinetic energy. When heated sufficiently, the mixture compresses and reduces in volume at such an extreme rate that a supersonic flow is produced within the combustion chamber.

This flame accelerates through impeller nozzles and pushes out against atmospheric pressure to generate thrusts force – around 41000N of static thrust for typical ramjet engines when powered by JP-7 jet fuel

2.Describe the construction and characteristics of different types of gas turbine engines.

The construction and characteristics of different types of gas turbine engines are as follows:

The gas generator cycle engine uses a high-pressure compressor to compress atmospheric air and then burns the compressed mixture with fuel in a combustion chamber. The hot gases generated are expanded through a turbine, which drives the compressor as well as an output shaft that can be used to power planes. This is the most common type of jet currently in use.

The turbojet uses the gas generator cycle; it was developed as a simpler variation of Frank Whittle’s original design in which all components were mixed together, thus requiring lesser mechanical complexity. Today turbofans account for most new engine deliveries due to their higher efficiency.

A turbofan differs from a turbojet by having larger fan ducts and adding an extra turbine stage to the basic turbojet. This leads to a significantly higher efficiency due mainly to reduced exhaust velocity and better thermal utilization of the cold expanded exhaust gas from the engine core.

Turbofans are also almost always classified as high-bypass jet engines, meaning that a large percentage of the air (as opposed to gasoline or diesel) entering the front of the engine bypasses the combustion chamber and nozzle, giving a much greater thrust force for a given amount of fuel consumed.

Modern turbofans have very high mass flow rates of air when compared with subsonic types; this gives them much greater thrust per unit weight at takeoff, making them suitable for use on large aircraft in which low fuel consumption is desirable.

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3.Illustrate the performance of aerospace propulsion systems.

The performance of aerospace propulsion systems can be illustrated by comparing the specific impulse with the fuel efficiency of a typical passenger car.

The specific impulse is defined as the thrust (force) per unit mass flow rate and is represented in SI units by seconds (the symbol for which is s). It can also be measured in slugs/ft³ or pound-force seconds/LBM.

It should not be confused with momentum; an increase in the number of seconds implies an increase in momentum but does not necessitate a change in kinetic energy and vice versa. The effective exhaust velocity, v e, often differs from this figure because it accounts for losses such as atmospheric drag, rocket nozzle efficiency, and heat transfer to ambient air.

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The theoretical maximum value of impulse is limited to the change in kinetic energy of a rocket, delta K e. When this occurs in one dimension, then impulse is equal to the change in momentum of the vehicle (which equals Delta T times mass).

For rockets operating in several (3 or more) dimensions, taking into account all relevant parameters defines the total impulse and maximum effective exhaust velocity.

For example, a theoretical particle accelerator might have many stages: i=1 gives momenta; i=2 gives velocities at each stage exit; I =3 calculates ∆K for each step in acceleration (to very high energies), etc.; then systematically takes into account losses from non-ideal operation as functions of force (i.e., fixed efficiency x actual thrust), etc. Total impulse can be obtained and is the product of I and time (i.e., delta K over delta t).

For rockets operating in only one dimension (such as a ground-level rocket sled to measure vehicle performance), the extra complexity needed to account for all relevant parameters defining total impulse might not be warranted. The increase in exhaust velocity may not be worth the effort required for its definition.

4.Analyze individual component performance and component matching

We can analyze individual component performance and component matching by calculating thermal efficiency. Thermal efficiency is the ratio of output to input energy, or The total heat energy put into a rocket engine by its oxidizer and fuel must equal the kinetic and potential energy created by their burning in reaction mass JxD (i.e., where D = propellant density). Theoretical two-stage vehicles often have three component areas:

If single-component performance is known, then a theoretical versus experimental comparison can be made. Alternatively, if individual components are not available (or are otherwise unavailable), then an overall systematic approach can be applied towards calculating theoretical performance.

5.Discuss the methodologies used in improving efficiencies and increasing specific work outputs of an aerospace propulsion system.

The methodologies used in improving efficiencies and increasing specific work outputs of an aerospace propulsion system are the following:

Developing improved hardware such as new and better materials used in the turbine, compressor, or diffuser.

The turbines of aerospace engines are designed for high performance, lightweight, and low cost. The flow area is large (up to 100%) at all stages because large mass flows allow shorter blades with a lower tip speed ratio (better energy efficiency). They are usually very heavy but lightweight is more important than mass in an aerospace application where they may be subject to vibration loads of 1000–4000 g and thermal gradients up to 800°C.

6.Solve a Parametric Cycle Analysis for Ideal Engines.

We can solve a parametric cycle analysis for ideal engines by using the information outlined above. The inputs for this analysis are:

Total pressure ratio,  “P 2 / P 1 ” =  (1427/990) =  1.5085

When solving for the compression ratio and turbine inlet temperature using the above equations we have to consider that there is a perfect gas and a real gas assumption made with each equation.

In terms of calculating the compressor inlet temperature, we really need to consider the actual heat capacity of air as well as any heat transfer due to friction between blades and casing when calculating T 2 . With these considerations in mind let’s solve for a specific value of γ. For this calculation, I will use Cp = 0.2468 kJ/kg -K and cv = 0.710 kJ/kg K for air at 120 °F (49 °C). Also, consider an air flow rate of 100lb/sec = SFM =  10 volumetric units.

For a realistic value of γ, I recommend using the average between 1.4 and 6.5 kg/m-K which would be about 4 kg/ m-K.”

The above is old Tom Myron information and not entirely accurate with modern gas turbines which use compression heating methods to increase processor temperatures from T2 to T3.

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7.Calculate a Parametric Cycle Analysis for Real Engines.

We can calculate a parametric cycle analysis for real engines. The mathematical approach is to calculate the engine efficiency factor (η) at two points, T1 and T2. Then, based on some arbitrary mass flow rate (M), we can solve for the speed of a compressor in RPMs needed to achieve that specific mass flow rate. We will have two equations with two unknowns. The density ratio and RPM ratio as functions of gas pressure ratio must be calculated beforehand using our compressible chart in order to make our parametric cycle work correctly.

There are many important factors affecting compressor tip speed but for now, I will only list those factors affecting propelling aircraft through the air:

  1. Altitude
  2. Ambient Pressure Ratio Effect
  3. Mass Flow Rate
  4. RPM
  5. Efficiency of Compressor (NPSHR)
  6. Gas Composition & Pressure Ratio
  7. Density Ratio (D/P)
  8. Design of Compressor and HP Turbine Are the RPMs fixed?

The design specifications for a compressor are not only its speed but also pressure ratio, mass flow rate as well as inlet conditions such as temperature and humidity to name a few (see this article for more information). This means that propelling an aircraft through the air is not only a matter of changing a parameter such as the RPMs in our example above, but it will involve performing trade-off analysis between those parameters themselves and other design factors affecting the overall efficiency of the engine.

8.Apply self-study and individual lab assignment skills learned on the course in aerospace MRO work.

We apply self-study and individual lab assignment skills learned on the course in aerospace MRO work, in the following ways.

Using individual lab assignment skills, each graduate is required to convert or re-design a specific part. We have examples of landing gear actuator and engine nacelle actuator designs that have been completed by our graduates for aerospace MRO work at companies such as GKN Aerospace (formerly Fokker).

These are self-study projects where students look at the current design of the part and evaluate numerous methods to carry out high-efficiency transformations of them. In addition, they must determine if any modifications need to be made in order for their new design to function correctly within its original application.

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