DOCSLIB.ORG

Understanding the Rocket Engine Performance in BLOODHOUND SSC the BLOODHOUND Engineering Project Mechanical Engineering

Total Page：16

File Type：pdf, Size：1020Kb

Download full-text PDF Read full-text
Understanding the Rocket Engine Performance in BLOODHOUND SSC the BLOODHOUND Engineering Project Mechanical Engineering Understanding the Rocket Engine Performance in BLOODHOUND SSC The BLOODHOUND Engineering Project Mechanical Engineering This exemplar requires some knowledge of Physics and Chemistry together with Maths. INTRODUCTION SCENARIO Figure 1: The BLOODHOUND SSC The BLOODHOUND Project is an iconic engineering and education adventure that is pushing technology to its limit, providing us with a once in a lifetime opportunity to inspire the next generation of scientists and engineers. BLOODHOUND SSC (Super Sonic Car) has been designed to set a new land speed record of 1,000 miles per hour (mph). BLOODHOUND SSC is powered by a EUROJET EJ200 jet engine and a hybrid rocket engine. A rocket is basically an engine which carries both a fuel and oxidiser (it does not require any Figure 2: Static tests of the 15.2cm (6-inch) hybrid oxygen from the air). Hybrid rockets use a solid chamber fuel and a liquid oxidiser. The fuel is contained within the combustion chamber and the liquid The two pictures above in figure 2 show static oxidiser is injected at the top of the chamber. tests of the 15.2cm (6-inch) diameter hybrid Therefore a hybrid can be easily shutdown by rocket chamber that is being used to develop the turning off the supply of liquid oxidiser, making hybrid rocket for BLOODHOUND SSC. them more suitable for use in a land speed The hybrid rocket in BLOODHOUND SSC uses record car. Hybrids use the same oxidisers as HTP (a concentrated hydrogen peroxide H2O2, liquid propellant systems; fuels can include basically water with an extra oxygen atom) as synthetic rubbers and plastics (such as PVC). the oxidiser and a synthetic rubber Hydroxyl- IMPORTANCE OF EXEMPLAR IN REAL LIFE Terminated Polybutadiene (HTPB) as the primary fuel. The hybrid combustion chamber for Thrust is a reaction force described BLOODHOUND SSC has a catalyst pack to quantitatively by Newton's Second and Third decompose the HTP, the decomposition Laws. When a system expels or accelerates products then enter the fuel grain. The fuel mass in one direction, the accelerated mass will automatically ignites in contact with the exert an equal but opposite force on that system. decomposition products, generating pressure. A rocket engine produces thrust by expelling hot The combustion products then enter the nozzle gas at a very high velocity. which converts the high pressure, low velocity Calculating thrust and specific impulse are gas into high velocity low pressure gas fundamental to the design of any rocket engine. generating up to 122 kN (27,500 lbs) of thrust. Achieving the optimum oxidiser:fuel (O:F) ratio The chamber contains 181 kg of fuel which can and understanding how the chamber pressure run for up to 20 seconds. The chamber is 45.7cm and the combustion process affect specific in diameter and 4.27m long. Development work impulse is the key to achieving an efficient has been conducted on a 15.2 cm (6-inch) system. diameter chamber. Several firings have been conducted to test various aspects of chamber 1 design, fuel grain configuration and catalyst The second expression in equation (2) shows material. that the thrust produced by a rocket varies with altitude because of the dependence on ambient One of the challenges is to find optimum specific pressure. Thus, rocket performance is usually impulse of the rocket while reducing the defined by two limits T (thrust at sea level) and harshness of the environment (in terms of high sl heat fluxes) experienced by key components of Tv (thrust in vacuum). The customary index of the motor (specifically the nozzle and throat). performance is specific impulse, Isp defined by: Calculating the heat transfer rate to rocket motor T components is frequently dealt with using CFD I = … (3) analysis and is not discussed here. Also, there sp dm g × are so many dynamic variables in a rocket dt combustion chamber that we need to simplify the where g is the acceleration due to gravity which equations in order to get an approximate mathematical answer. To try and calculate all the varies with altitude in the same way that T does. dynamic variables requires extremely complex Using specific impulse to characterise the rocket mathematics. performance is analogous to using miles per gallon or specific horsepower to characterise a Therefore, in this exemplar, we will look at some car. of the fundamental aspects of calculating rocket performance and examine some of the more In order to address the question of O:F, we need complex aspects which are key to achieving an to keep some other parameters constant. The efficient hybrid rocket for BLOODHOUND SSC. mathematical analysis is simplest if we assume that the motor designers expand the rocket MATHEMATICAL MODEL engine nozzle in such a way that the static A rocket is propelled forward by a thrust force pressure at exit ( Pe ) closely matches the static equal in magnitude, but opposite in direction, to pressure at the operating altitude ( P ). Hence, the time-rate of momentum change of the a exhaust gas accelerated from the combustion the second term in equation (2) is usually small. chamber through the rocket engine nozzle Thus, ignoring that term, we get: (shown in figure 2). dm T = v … (4) dt e Mathematically, it can be expressed as: Combining equation (3) and (4), we get: dm v T = v … (1) I = e or I ∝ v … (5) dt sp g sp e where This shows that Isp is strongly dependent on the T = Thrust exit velocity ve which can be calculated from the dm = Mass flow rate of exhaust following expression: dt ⎡ ()γ −1 ⎤ Speed of the exhaust gases measured 2γ ⎛ P ⎞ γ v = v RT ⎢1 ⎜ e ⎟ ⎥ relative to the rocket e = c ⎢ − ⎜ ⎟ ⎥ … (6) ()γ −1 ⎝ Pc ⎠ Please note that equation (1) is true only for a ⎣⎢ ⎦⎥ perfectly expanded nozzle. where In case of BLOODHOUND SSC, this thrust is γ = Ratio of specific heats expressed as follows: R = Specific gas constant dm T = v + ()P − P A … (2) T = Combustion chamber temperature dt e e a e c where Pc = Combustion chamber pressure (Please refer to the isentropic expansion explained in ve = Exit velocity of the exhaust gases Reference 3, 4 or 5 under the title “Where to find more” in order to get better understanding of equation (6) above. Pe = Rocket engine nozzle exit pressure Proof of this equation is beyond the scope of this exemplar.) Pa = Ambient pressure For a particular fuel:oxidiser combination, γ is a A = Exit plane area slowly varying function of O:F and hence can be e considered as constant for simplicity. The ratio Pe Pc can be considered as a constant here to 2 get a fair approximation of the specific impulse. combustibles are here by design to enhance Isp , However, variable ratio P P requires more e c i.e. they should not be regarded as unburnt fuel. complicated analysis and is not covered here. Isp is the main driver of hybrid rocket design Thus, from equation (6), we see that: rather than combustion efficiency. Hybrid rockets are often run oxidiser rich (O:F larger than ve ∝ RTc … (7) stoichiometric) since a large flow rate of oxidiser We know that the specific gas constant R is is required to liberate fuel vapour from the fuel defined as the ratio of the universal gas constant grain. As a consequence, the Isp from a hybrid is R and the mean molecular weight , i.e.: u Mm usually lower than a high efficiency liquid engine R (cryogenic H2/LOX). R = u Mm CONCLUSION Hence, In this exemplar, we have looked at fundamental rocket performance parameters such as 1 R ∝ … (8) calculating thrust and specific impulse and also Mm made an attempt to explain why rocket exhaust plumes frequently afterburn upon mixing with the where Mm is the mean molecular weight of the ambient air producing bright flames in the combustion products in this case. exhaust gas. This is because optimum specific Combining equation (7) and (8), we can say that: impulse is frequently achieved when light flammable elements remain present in the Tc exhaust plume i.e. not all the fuel is burnt down ve ∝ … (9) to its final heavier state (usually H2O or CO2). Mm WHERE TO FIND MORE Finally, observing equation (5) and (9) together, it is clear that the optimum I can only be 1. Basic Engineering Mathematics, John Bird, sp 2007, published by Elsevier Ltd. achieved when the ratio of the combustion 2. Engineering Mathematics, Fifth Edition, John temperature Tc to the mean molecular weight Bird, 2007, published by Elsevier Ltd. Mm of the combustion products is maximised. In 3. Mechanics of Fluids, B. S. Massey, need particular, this reasoning explains why many more details. rockets produce afterburning plumes. At first 4. Rocket Exhaust Plume Phenomenology, glance, the combustion of H2 or CO in the plume Frederick S. Simmons, 2000, Published by would appear to point the inefficiency of the the AIAA rocket engine since burning these materials within the rocket engine would result in the 5. Rocket Propulsion Elements, George P. release of more chemical energy. However, the Sutton , Donald M. Ross, Oscar Biblarz; 7th mean molecular weight of a hydrogen oxygen Edition mixture is considerably lower than for pure water 6. http://www.bloodhoundssc.com vapour. As such, we find that an optimum Isp is 7. http://moodle.student.cnwl.ac.uk/moodledata achieved at a value of O:F (oxidiser to fuel ratio) _shared/CDX%20eTextbook/dswmedia/fuelS that is significantly lower than the stoichiometric ys/gasoline/fund/stoichiometricratio.html ratio (please refer to Reference 7) for complete combustion. This leads to the presence of combustible products in the plume.
Load more

Recommended publications
  • Different Types of Rocket Nozzles
    Different Types of Rocket Nozzles 5322- Rocket Propulsion Project Report By Patel Harinkumar Rajendrabhai(1001150586) 1. Introduction 1.1 What is Nozzle and why they are used? A nozzle is a device designed to control the direction or characteristics of a fluid flow (especially to increase velocity) as it exits (or enters) an enclosed chamber or Pipe[9]. Nozzles are frequently used to control the rate of flow, speed, direction, mass, shape, and/or the pressure of the stream that emerges from them. In nozzle velocity of fluid increases on the expense of its pressure energy. A Water Nozzle[9] Rotator Style Pivot Sprinkler[9] 1.2 What is Rocket Nozzle? A rocket engine nozzle is a propelling nozzle (usually of the de Laval type) used in a rocket engine to expand and accelerate the hot gases from combustion so as to produce thrust according to Newton’s law of motion. Combustion gases are produced by burning the propellants in combustor, they exit the nozzle at very high Speed (hypersonic). 1.3 Properties of Rocket Nozzle Nozzle produces thrust. Exhaust gases from combustion are pushed into throat region of nozzle. Throat is smaller cross-sectional area than rest of engine, gases are compressed to high pressure. Nozzle gradually increases in cross-sectional area allowing gases to expand and push against walls creating thrust. Convert thermal energy of hot chamber gases into kinetic energy and direct that energy along nozzle axis.[1] Mathematically, ultimate purpose of nozzle is to expand gases as efficiently as possible so as to maximize exit velocity.[1] Rocket Engine[1] F m eVe Pe Pa Ae Neglecting Pressure losses F m eVe 2 Different types of Rocket Nozzle Configuration(shape) The rocket nozzles can have many shapes configurations.
    [Show full text]
  • Back to the the Future? 07> Probing the Kuiper Belt
    SpaceFlight A British Interplanetary Society publication Volume 62 No.7 July 2020 £5.25 SPACE PLANES: back to the the future? 07> Probing the Kuiper Belt 634089 The man behind the ISS 770038 Remembering Dr Fred Singer 9 CONTENTS Features 16 Multiple stations pledge We look at a critical assessment of the way science is conducted at the International Space Station and finds it wanting. 18 The man behind the ISS 16 The Editor reflects on the life of recently Letter from the Editor deceased Jim Beggs, the NASA Administrator for whom the building of the ISS was his We are particularly pleased this supreme achievement. month to have two features which cover the spectrum of 22 Why don’t we just wing it? astronautical activities. Nick Spall Nick Spall FBIS examines the balance between gives us his critical assessment of winged lifting vehicles and semi-ballistic both winged and blunt-body re-entry vehicles for human space capsules, arguing that the former have been flight and Alan Stern reports on his grossly overlooked. research at the very edge of the 26 Parallels with Apollo 18 connected solar system – the Kuiper Belt. David Baker looks beyond the initial return to the We think of the internet and Moon by astronauts and examines the plan for a how it helps us communicate and sustained presence on the lunar surface. stay in touch, especially in these times of difficulty. But the fact that 28 Probing further in the Kuiper Belt in less than a lifetime we have Alan Stern provides another update on the gone from a tiny bleeping ball in pioneering work of New Horizons.
    [Show full text]
  • Aerospace Engine Data
    AEROSPACE ENGINE DATA Data for some concrete aerospace engines and their craft ................................................................................. 1 Data on rocket-engine types and comparison with large turbofans ................................................................... 1 Data on some large airliner engines ................................................................................................................... 2 Data on other aircraft engines and manufacturers .......................................................................................... 3 In this Appendix common to Aircraft propulsion and Space propulsion, data for thrust, weight, and specific fuel consumption, are presented for some different types of engines (Table 1), with some values of specific impulse and exit speed (Table 2), a plot of Mach number and specific impulse characteristic of different engine types (Fig. 1), and detailed characteristics of some modern turbofan engines, used in large airplanes (Table 3). DATA FOR SOME CONCRETE AEROSPACE ENGINES AND THEIR CRAFT Table 1. Thrust to weight ratio (F/W), for engines and their crafts, at take-off*, specific fuel consumption (TSFC), and initial and final mass of craft (intermediate values appear in [kN] when forces, and in tonnes [t] when masses). Engine Engine TSFC Whole craft Whole craft Whole craft mass, type thrust/weight (g/s)/kN type thrust/weight mini/mfin Trent 900 350/63=5.5 15.5 A380 4×350/5600=0.25 560/330=1.8 cruise 90/63=1.4 cruise 4×90/5000=0.1 CFM56-5A 110/23=4.8 16
    [Show full text]
  • Development of Turbopump for LE-9 Engine
    Development of Turbopump for LE-9 Engine MIZUNO Tsutomu : P. E. Jp, Manager, Research & Engineering Development, Aero Engine, Space & Defense Business Area OGUCHI Hideo : Manager, Space Development Department, Aero Engine, Space & Defense Business Area NIIYAMA Kazuki : Ph. D., Manager, Space Development Department, Aero Engine, Space & Defense Business Area SHIMIYA Noriyuki : Space Development Department, Aero Engine, Space & Defense Business Area LE-9 is a new cryogenic booster engine with high performance, high reliability, and low cost, which is designed for H3 Rocket. It will be the first booster engine in the world with an expander bleed cycle. In the designing process, the performance requirements of the turbopump and other components can be concurrently evaluated by the mathematical model of the total engine system including evaluation with the simulated performance characteristic model of turbopump. This paper reports the design requirements of the LE-9 turbopump and their latest development status. Liquid oxygen 1. Introduction turbopump Liquid hydrogen The H3 rocket, intended to reduce cost and improve turbopump reliability with respect to the H-II A/B rockets currently in operation, is under development toward the launch of the first H3 test rocket in FY 2020. In rocket development, engine is an important factor determining reliability, cost, and performance, and as a new engine for the H3 rocket first stage, an LE-9 engine(1) is under development. A rocket engine uses a turbopump to raise the pressure of low-pressure propellant supplied from a tank, injects the pressurized propellant through an injector into a combustion chamber to combust it under high-temperature and high- pressure conditions.
    [Show full text]
  • 6. Chemical-Nuclear Propulsion MAE 342 2016
    2/12/20 Chemical/Nuclear Propulsion Space System Design, MAE 342, Princeton University Robert Stengel • Thermal rockets • Performance parameters • Propellants and propellant storage Copyright 2016 by Robert Stengel. All rights reserved. For educational use only. http://www.princeton.edu/~stengel/MAE342.html 1 1 Chemical (Thermal) Rockets • Liquid/Gas Propellant –Monopropellant • Cold gas • Catalytic decomposition –Bipropellant • Separate oxidizer and fuel • Hypergolic (spontaneous) • Solid Propellant ignition –Mixed oxidizer and fuel • External ignition –External ignition • Storage –Burn to completion – Ambient temperature and pressure • Hybrid Propellant – Cryogenic –Liquid oxidizer, solid fuel – Pressurized tank –Throttlable –Throttlable –Start/stop cycling –Start/stop cycling 2 2 1 2/12/20 Cold Gas Thruster (used with inert gas) Moog Divert/Attitude Thruster and Valve 3 3 Monopropellant Hydrazine Thruster Aerojet Rocketdyne • Catalytic decomposition produces thrust • Reliable • Low performance • Toxic 4 4 2 2/12/20 Bi-Propellant Rocket Motor Thrust / Motor Weight ~ 70:1 5 5 Hypergolic, Storable Liquid- Propellant Thruster Titan 2 • Spontaneous combustion • Reliable • Corrosive, toxic 6 6 3 2/12/20 Pressure-Fed and Turbopump Engine Cycles Pressure-Fed Gas-Generator Rocket Rocket Cycle Cycle, with Nozzle Cooling 7 7 Staged Combustion Engine Cycles Staged Combustion Full-Flow Staged Rocket Cycle Combustion Rocket Cycle 8 8 4 2/12/20 German V-2 Rocket Motor, Fuel Injectors, and Turbopump 9 9 Combustion Chamber Injectors 10 10 5 2/12/20
    [Show full text]
  • Combustion Tap-Off Cycle
    College of Engineering Honors Program 12-10-2016 Combustion Tap-Off Cycle Nicole Shriver Embry-Riddle Aeronautical University, [email protected] Follow this and additional works at: https://commons.erau.edu/pr-honors-coe Part of the Aeronautical Vehicles Commons, Other Aerospace Engineering Commons, Propulsion and Power Commons, and the Space Vehicles Commons Scholarly Commons Citation Shriver, N. (2016). Combustion Tap-Off Cycle. , (). Retrieved from https://commons.erau.edu/pr-honors- coe/6 This Article is brought to you for free and open access by the Honors Program at Scholarly Commons. It has been accepted for inclusion in College of Engineering by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. Honors Directed Study: Combustion Tap-Off Cycle Date of Submission: December 10, 2016 by Nicole Shriver [email protected] Submitted to Dr. Michael Fabian Department of Aerospace Engineering College of Engineering In Partial Fulfillment Of the Requirements Of Honors Directed Study Fall 2016 1 1.0 INTRODUCTION The combustion tap-off cycle is also known as the “topping cycle” or “chamber bleed cycle.” It is an open liquid bipropellant cycle, usually of liquid hydrogen and liquid oxygen, that combines the fuel and oxidizer in the main combustion chamber. Gases from the edges of the combustion chamber are used to power the engine’s turbine and are expelled as exhaust. Figure 1.1 below shows a picture representation of the cycle. Figure 1.1: Combustion Tap-Off Cycle The combustion tap-off cycle is rather unconventional for rocket engines as it has only been put into practice with two engines.
    [Show full text]
  • Materials for Liquid Propulsion Systems
    https://ntrs.nasa.gov/search.jsp?R=20160008869 2019-08-29T17:47:59+00:00Z CHAPTER 12 Materials for Liquid Propulsion Systems John A. Halchak Consultant, Los Angeles, California James L. Cannon NASA Marshall Space Flight Center, Huntsville, Alabama Corey Brown Aerojet-Rocketdyne, West Palm Beach, Florida 12.1 Introduction Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks which provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de-Laval nozzle) which forces them to accelerate; then as the nozzle flares outwards, they expand and further accelerate. It is the mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, which produce thrust according to Newton’s third law: for every action there is an equal and opposite reaction. [1] Solid rocket motors are cheaper to manufacture and offer good values for their cost.
    [Show full text]
  • Enhancement of Volumetric Specific Impulse in HTPB/Ammonium Nitrate Mixed
    Utah State University DigitalCommons@USU All Graduate Plan B and other Reports Graduate Studies 12-2016 Enhancement of Volumetric Specific Impulse in TPB/H Ammonium Nitrate Mixed Hybrid Rocket Systems Jacob Ward Forsyth Utah State University Follow this and additional works at: https://digitalcommons.usu.edu/gradreports Part of the Propulsion and Power Commons Recommended Citation Forsyth, Jacob Ward, "Enhancement of Volumetric Specific Impulse in TPB/AmmoniumH Nitrate Mixed Hybrid Rocket Systems" (2016). All Graduate Plan B and other Reports. 876. https://digitalcommons.usu.edu/gradreports/876 This Report is brought to you for free and open access by the Graduate Studies at DigitalCommons@USU. It has been accepted for inclusion in All Graduate Plan B and other Reports by an authorized administrator of DigitalCommons@USU. For more information, please contact [email protected]. ENHANCEMENT OF VOLUMETRIC SPECIFIC IMPULSE IN HTPB/AMMONIUM NITRATE MIXED HYBRID ROCKET SYSTEMS by Jacob W. Forsyth A report submitted in partial fulfillment of the requirements for the degree of MASTER OF SCIENCE in Aerospace Engineering Approved: ______________________ ____________________ Stephen A. Whitmore Ph.D. David Geller Ph.D. Major Professor Committee Member ______________________ Rees Fullmer Ph.D. Committee Member UTAH STATE UNIVERSITY Logan, Utah 2016 ii Copyright © Jacob W. Forsyth 2016 All Rights Reserved iii ABSTRACT Enhancement of Volumetric Specific Impulse in HTPB/Ammonium Nitrate Mixed Hybrid Rocket Systems by Jacob W. Forsyth, Master of Science Utah State University, 2016 Major Professor: Dr. Stephen A. Whitmore Department: Mechanical and Aerospace Engineering Hybrid rocket systems are safer and have higher specific impulse than solid rockets. However, due to large oxidizer tanks and low regression rates, hybrid rockets have low volumetric efficiency and very long longitudinal profiles, which limit many of the applications for which hybrids can be used.
    [Show full text]
  • Validation of a Simplified Model for Liquid Propellant Rocket Engine Combustion Chamber Design
    IOP Conference Series: Materials Science and Engineering PAPER • OPEN ACCESS Validation of a simplified model for liquid propellant rocket engine combustion chamber design To cite this article: M Hegazy et al 2020 IOP Conf. Ser.: Mater. Sci. Eng. 973 012003 View the article online for updates and enhancements. This content was downloaded from IP address 170.106.33.14 on 25/09/2021 at 23:25 AMME-19 IOP Publishing IOP Conf. Series: Materials Science and Engineering 973 (2020) 012003 doi:10.1088/1757-899X/973/1/012003 Validation of a simplified model for liquid propellant rocket engine combustion chamber design M Hegazy1, H Belal2, A Makled3 and M A Al-Sanabawy4 1 M.Sc. Student, Rocket Department, Military Technical College, Egypt 2 Assistant Professor, Rocket Department, Military Technical College, Egypt 3 Associate Professor. Zagazig University, Egypt 4 Associate Professor. Rocket Department, Military Technical College, Egypt [email protected] Abstract. The combustion phenomena inside the thrust chamber of the liquid propellant rocket engine are very complicated because of different paths for elementary processes. In this paper, the characteristic length (L*) approach for the combustion chamber design will be discussed compared to the effective length (Leff) approach. First, both methods are introduced then applied for real LPRE. The effective length methodology is introduced starting from the basic model until developing the empirical equations that may be used in the design process. The classical procedure of L* was found to over-estimate the required cylindrical length in addition to the inherent shortcoming of not giving insight where to move to enhance the design.
    [Show full text]
  • Helicopter Turboshafts
    Helicopter Turboshafts Luke Stuyvenberg University of Colorado at Boulder Department of Aerospace Engineering The application of gas turbine engines in helicopters is discussed. The work- ings of turboshafts and the history of their use in helicopters is briefly described. Ideal cycle analyses of the Boeing 502-14 and of the General Electric T64 turboshaft engine are performed. I. Introduction to Turboshafts Turboshafts are an adaptation of gas turbine technology in which the principle output is shaft power from the expansion of hot gas through the turbine, rather than thrust from the exhaust of these gases. They have found a wide variety of applications ranging from air compression to auxiliary power generation to racing boat propulsion and more. This paper, however, will focus primarily on the application of turboshaft technology to providing main power for helicopters, to achieve extended vertical flight. II. Relationship to Turbojets As a variation of the gas turbine, turboshafts are very similar to turbojets. The operating principle is identical: atmospheric gases are ingested at the inlet, compressed, mixed with fuel and combusted, then expanded through a turbine which powers the compressor. There are two key diferences which separate turboshafts from turbojets, however. Figure 1. Basic Turboshaft Operation Note the absence of a mechanical connection between the HPT and LPT. An ideal turboshaft extracts with the HPT only the power necessary to turn the compressor, and with the LPT all remaining power from the expansion process. 1 of 10 American Institute of Aeronautics and Astronautics A. Emphasis on Shaft Power Unlike turbojets, the primary purpose of which is to produce thrust from the expanded gases, turboshafts are intended to extract shaft horsepower (shp).
    [Show full text]
  • Session 6:Analytical Approximations for Low Thrust Maneuvers
    Session 6: Analytical Approximations for Low Thrust Maneuvers As mentioned in the previous lecture, solving non-Keplerian problems in general requires the use of perturbation methods and many are only solvable through numerical integration. However, there are a few examples of low-thrust space propulsion maneuvers for which we can find approximate analytical expressions. In this lecture, we explore a couple of these maneuvers, both of which are useful because of their precision and practical value: i) climb or descent from a circular orbit with continuous thrust, and ii) in-orbit repositioning, or walking. Spiral Climb/Descent We start by writing the equations of motion in polar coordinates, 2 d2r �dθ � µ − r + = a (1) 2 2 r dt dt r 2 d θ 2 dr dθ aθ + = (2) 2 dt r dt dt r We assume continuous thrust in the angular direction, therefore ar = 0. If the acceleration force along θ is small, then we can safely assume the orbit will remain nearly circular and the semi-major axis will be just slightly different after one orbital period. Of course, small and slightly are vague words. To make the analysis rigorous, we need to be more precise. Let us say that for this approximation to be valid, the angular acceleration has to be much smaller than the corresponding centrifugal or gravitational forces (the last two terms in the LHS of Eq. (1)) and that the radial acceleration (the first term in the LHS in the same equation) is negligible. Given these assumptions, from Eq. (1), dθ r µ d2θ 3 r µ dr ≈ ! ≈ − (3) 3 2 5 dt r dt 2 r dt Substituting into Eq.
    [Show full text]
  • Thrust Reverser
    DC-10THRUST REVERSER Following an airplane accident in which inadver- tent thrust reverser deployment was considered a major contributor, the aviation industry and SAFETY the U.S. Federal Aviation Administration (FAA) adopted new criteria for evaluating the safety of ENHANCEMENT thrust reverser systems on commercial airplanes. Several airplane models were determined to be uncontrollable in some portions of the flight HARRY SLUSHER envelope after inadvertent deployment of the PROGRAM MANAGER AIRCRAFT MODIFICATION ENGINEERING thrust reverser. In response, Boeing and the BOEING COMMERCIAL AIRPLANES GROUP SAFETY FAA issued service bulletins and airworthiness AERO directives, respectively, for mandatory inspec- 39 tions and installation of thrust reverser actuation system locks on affected Boeing-designed airplanes. Boeing and the FAA are issuing similar documents for all models of the DC-10. air seal, fairing, and the aft frame; MODIFICATION OF THE THRUST a proposed AD for the indication reverser. This allows operators to stock and checks of the feedback rod-to-yoke 2 system on all models of the DC-10. neutral spare reversers and quickly The safety enhancement program REVERSER INDICATION SYSTEM oeing has initiated a four- alignment, translating cowl auto- An AD mandating the incorporation of configure them for any engine position. phased safety enhancement for the DC-10 is designed to improve restow function, position indication The second phase of the DC-10 safety B enhancement program involves modify- Boeing Service Bulletin DC10-78-060 The basic design elements of the program for thrust reverser the reliability of the thrust reverser for the overpressure shutoff valve, is expected by the end of 2000.
    [Show full text]