The Top Athletes Looking for an Edge and the Scien Essays - Sports

The Top Athletes Looking for an Edge and the Scientists Trying to Catch Them. Behind the scenes there will be a high-tech, high-stakes competition between Olympic athletes who use banned substances and drug testers out to catch them ByChristie Aschwanden Smithsonian Magazine | Subscribe July 2012 D eeDee Trotter was on an airplane in 2006 when she overheard a passenger seated behind her discussing the steroids scandal. Federal investigators in the Balco case, named for a lab that produced supplements, would eventually implicate more than two dozen athletes for the use of performance-enhancing drugs, including Barry Bonds, baseball's home run king, and Marion Jones, the track-and-field star, who would end up in jail, stripped of five Olympic medals. "This guy was reading the newspaper and he said, Oh, they're all on drugs,'" recalls Trotter, a runner who won a gold medal in the 4 x 400 meter relay at the 2004 Olympics. She was furious. "I turned around and said, Heyexcuse me, I'm sorry, but that's not true. I'm a professional athlete and Olympic gold medalist, and I'm not on drugs. I've never even considered it. ' " Currently vying to join the U.S. team and appear in her third Olympics, Trotter projects a sassy confidence. "It really upset me that it's perceived that waythat if she runs fast, then she's on drugs. I hated that and I gave him a little attitude." That airplane conversation prompted Trotter to create a foundation called Test Me, I'm Clean! "It gave us clean athletes a chance to defend ourselves," says Trotter. "If you see someone wearing this wristband"she holds up a rubbery white bracelet emblazoned with the group's name " it means that I am a clean athlete. I do this with hard work, honesty and honor. I don't take any outside substances." As Trotter tells me this story, I catch myself wondering if it's all just a bunch of pre-emptive PR. It pains me to react this way, but with doping scandals plaguing the past three Summer Olympics and nearly every disgraced athlete insisting, at least initially, that he or she is innocent, it's hard to take such protestations at face value. My most profound disillusionment came from a one-time friend, Tyler Hamilton, my teammate on the University of Colorado cycling team. When he won a gold medal in the time trial at the 2004 Olympics, I was thrilled to see someone I'd admired as honest and hardworking reach the top of a sport that had been plagued by doping scandals. But in the days that followed, a new test implicated Hamilton for blood doping. His supporters began hawking "I Believe Tyler" T-shirts, and he took donations from fans to fund his defense. The evidence against him seemed indisputable, but the Tyler I knew in college was not a cheat or liar. So I asked him straight-out if he was guilty. He looked me in the eye and told me he didn't do it. Last year, after being subpoenaed by federal investigators, Hamilton finally confessed and returned his medal. The downfall of Olympic heroes has cast a cloud of suspicion over sports. And the dopers' victims aren't just the rivals from whom they stole their golden podium moments but every clean athlete whose performance is greeted with skepticism. Doping, or using a substance to enhance performance, is nothing new. Contrary to romantic notions about the purity of Olympic sports, ancient Greeks ingested special drinks and potions to give them an edge, and at the 1904 Games, athletes downed potent mixtures of cocaine, heroin and strych - nine. For most of Olympic history, using drugs wasn't considered cheating. Then, in the 1960 Olympics, Danish cyclist Knut Jensen passed out during a race, cracked his skull and later died. The coroner blamed the death on amphetamines, and the case led to anti-doping rules. Drug testing began with the 1968 Games, with a goal to protect athlete health. In addition to short-term damage, certain drugs also appear to increase the risk of heart disease and possibly cancer. The original intent of anti-doping rules was to prevent athletes from dropping dead of overdoses, but over the years the rules have come to focus just as intently on

Mid-Autumn Festival Essays - Autumn, Public Holidays In China

Mid-Autumn Festival Essays - Autumn, Public Holidays In China Mid-Autumn Festival The Mid-Autumn Festival occurs every year on the fifteenth day of the eighth month. This date is in respect to the lunar calendar which is used by the Chinese. In the Gregorian calendar, used in America, this day would be approximately the fifteenth of September. On this day, the moon is supposed to be at its fullest and brightest of the year. The whole family eats out or in their yards to celebrate and watch the full moon. Children play with paper lanterns and the same lanterns are hung outside the front doors of buildings, such as houses and restaurants. Mooncakes are eaten and Chinese tea is usually used to wash it down. The name, mooncake, is self-explanatory. It is a round cake, in the shape of a moon. The ingredients of the cake consist of lotus seeds, made into a sort of paste. The paste is surrounded by a crust, which usually has four Chinese characters imprinted on the top. These characters either tell the type of mooncake it is (i.e. regular, lotus with egg yolk), the name of the store it was bought from, or just simply says ?mooncake?. The origin of the mooncake is in China, during the Sang Dynasty. The Han people were conquered by the Mongolians and named the new dynasty Yuan. The Han people did not like living under Mongolian rule. Therefore, they wanted to rebel and retake China. However, the Mongolians had taken this into consideration and did not allow the people to communicate (especially public gatherings) or to possess sharp, pointed weaponry. Thus, the people had to find a way of communicating secretly. One group of men thought up the idea of placing a piece of paper with the date of the rebellion inside little cakes, which they would sell to the people, who would read the paper and find out the date. To gain permission from the Mongolian soldiers to sell the cakes, they told them that the cakes were a sort of offering to the gods. They said that they would pray that the Mongolian emperor could have eternal life. The gullible soldiers quickly agreed. Everyone received the cakes and the rebellion date was set for the fifteenth day of the eighth month. Since the Mongolians could not read Chinese, they did not know of the rebellion, were caught by surprise, and defeated. From then on, the fifteenth day of the eighth month was known as the day of the Mid-Autumn Festival to celebrate the day of the rebellion. Many myths are formed about holidays. One which goes with this holiday is about a time when the world had ten suns and the earth was hot and dry. Nothing could survive. A hero stepped forward and used nine arrows to shoot down nine of the suns. He was crowned king and married a beautiful wife. Within years of his reign, he became selfish and greedy, a dictator. He wanted to live forever and make the people suffer. Therefore, he mixed a powerful potion and made a pill which, when eaten, would give the person eternal life. His wife found him out and stole the pill. To keep her husband from eating it, she ate it herself. However, after she ate it, she felt her body get lighter and lighter until she was floating. She kept rising higher and higher until she reached the moon, where she lives until this day. There are many variations of this story, such as the bringing of a rabbit with her because the gods wanted to reward her bravery by giving her company for her loneliness. Some people say that they can sometimes see a woman in the moon with a rabbit and a tree (another variation).

Assignment B wk3 Essay Example | Topics and Well Written Essays - 750 words

Assignment B wk3 - Essay Example It is biblically that God is a God of truth and His reverence must be in observance with that truth. To worship refers to demonstrate honor and respect to God, and at what moment in his physical presence, that denotes to prostrate oneself in a way in which one demonstrates his supremacy over oneself (MacArthur, 1983). Worshipping in truth is to show adoration to him in human nature through the actions. The concept of praising God in spirit and truth generates from Jesus’ discussion with the lady at the well in John 4. In the discussion, the lady was talk about places of worship with Jesus, claiming that the Samaritans worshipped at Mount Gerizim while the Jews prayed at Jerusalem. Jesus had just demonstrated that He understood about her numerous spouses, also the fact that the present man she stayed with was not her spouse (MacArthur, 1983). This made her uneasy, so she tried to sidetrack His attention from her private life to issues of religion. Jesus did not get distracted from His session on right worship and got to the core of the issue when he said that the hour was coming, and it was already time, when the true worshipers would adore God in truth and spirit, for God needs such to worship Him in John 4. The overall message concerning worshipping God in spirit and truth is that adoration of the Father is not to be restricted to a solitary geographical site or essentially controlled by the temporary requirements of bible law (MacArthur, 1983). With the presence of Christ, the severance between Gentile and Jew was no longer pertinent, nor was the site of worship as the temple. With the Christ, all of God’s believers gained equivalent admission to God through Jesus. Worship became an issue of the spirit, not outside actions, and bound for by reality rather than ceremonial. True worship ought to be in spirit that

Psychology of religion Essay Example | Topics and Well Written Essays - 1250 words

Psychology of religion - Essay Example needs to acts in his own interests and to ward off things that will harm him.† (Retrieved from www.englishforums.com) Different communities maintain different religious beliefs, though attributes affiliated with the Supreme Being are similar to some extent. The same is the case with Christianity. Christianity is the most popular religion of the globe, as its followers are highest in number to all other faiths existing in the contemporary world. Christians have extracted the attributes of God from Biblical themes and stories. But the philosophers contain diversified opinion regarding the background, description, existence and executions of these characteristics. Hodgson & King (1985) have discussed the philosophical views of eminent theologian and philosopher Thomas Aquinas of 13th century with the work of contemporary thinker Gordon Kaufman, in a comparative way, in their famous work under the title â€Å"Readings In Christian Mythology.† The work concentrates on the religious aspects of Christianity with reference of religious themes and beliefs in order to show the relation of human characteristics with those attributed with Almighty God. St. Aquinas is of the view that the merit and demerit of all the qualities obtained and possessed by humans have been determined as good or bad by Almighty God. In other words, it is no t man that decides an action, an idea, a notion or a concept as good or bad; rather, these qualities have already been decided by the Lord, on the basis of which all the actions, activities and attitude of human beings are regulated and maintained. â€Å"All that man is, and can, and has†, Aquinas suggests, â€Å"must be referred to God; and therefore every action of man, whether good or bad, acquires merit or demerit in the sight of God, as far as the action itself is concerned.† (Quoted in Porter, 1997: 212) In the same way, Aquinas submits that no words in any language can portray the attributes of God. On the other hand, man has learnt and

The Relationship between Employee Commitment and Employee Engagement, Assignment

The Relationship between Employee Commitment and Employee Engagement, Employee Satisfaction - Assignment Example It also can be referred as creating a healthy work environment for the employees in order to motivate them. It will help the employees to connect with their work and job responsibilities (Storey, Wright and Ulrich, 2009, p.300). On the other hand, commitment can be defined as willingness to persevere in a course of reluctance and action to change plans. The employees devote their energy and time to fulfil their job responsibility as well as their personal, community, family and spiritual obligations. Employees, who are committed to their organizations and highly engaged in their job, provide effective competitive advantages to the organizations in terms of higher output. Uncommitted employees do not bother about workplace performance and outputs. On the other hand, the committed employees tend to provide their total effort to fulfil their personal career goals and job responsibility. Engagement of an employee cannot possible without effective commitment towards the organization and s eer hard work. Leaders or the managers of an organization play a vital role in employee engagement. It is important for a manager to provide value to the needs or satisfaction level of an employee in order to retrain employee commitment and employee engagement. Only a motivated employee can perform effectively in an organization. ... It will help an organization to achieve success (Mannelly, 2009, p.161). Committed employees are more engaged to their job and organization comparing to the uncommitted employees. Employee engagement, employer practices, work performance and business results are highly related to each other. It is the responsibility of the employers to motivate their employees to perform efficiently. Effective performance appraisal, incentive systems, career growth opportunities are the motivation and performance drivers for an employee in an organization. These aspects made an employee committed to their job. Committed employees provide their best performance in order to capitalize on the potential career opportunities. Therefore, it can be stated that, effective employee engagement can help an organization to increase its business productivity. Effective performance appraisal system increases the commitment level of an employee. It is evident that the global workplace behaviour is changing dramatic ally (Albrecht, 2010, p.67). Now-a-days, the customers are trying to achieve value added and high quality products and services. Therefore, the global organizations are trying to motivate their workforce in order to meet with the demand of the customers. The uncommitted employees cannot perform effectively due to lack of workplace motivation. As the skilled and motivated employees are the biggest assets of an organization, therefore it is responsibility of the organization to take care of their needs. Therefore, it can be concluded that committed employees are more engaged with their work and responsibilities than the uncommitted employees. Is it correct to say that Committed Employees are more satisfied than

Single Stage to Orbit (SSTO) Propulsion System

Single Stage to Orbit (SSTO) Propulsion System This paper discusses the relevant selection criteria for a single stage to orbit (SSTO) propulsion system and then reviews the characteristics of the typical engine types proposed for this role against these criteria. The engine types considered include Hydrogen/Oxygen (H2/O2) rockets, Scramjets, Turbojets, Turborockets and Liquid Air Cycle Engines. In the authors opinion none of the above engines are able to meet all the necessary criteria for an SSTO propulsion system simultaneously. However by selecting appropriate features from each it is possible to synthesise a new class of engines which are specifically optimised for the SSTO role. The resulting engines employ precooling of the airstream and a high internal pressure ratio to enable a relatively conventional high pressure rocket combustion chamber to be utilised in both airbreathing and rocket modes. This results in a significant mass saving with installation advantages which by careful design of the cycle thermodynamics enable s the full potential of airbreathing to be realised. The SABRE engine which powers the SKYLON launch vehicle is an example of one of these so called Precooled hybrid airbreathing rocket engines and the conceptual reasoning which leads to its main design parameters are described in the paper. Keywords: Reusable launchers, SABRE, SKYLON, SSTO 1.Introduction Several organisations world-wide are studying the technical and commercial feasibility of reusable SSTO launchers. This new class of vehicles appear to offer the tantalising prospect of greatly reduced recurring costs and increased reliability compared to existing expendable vehicles. However achieving this breakthrough is a difficult task since the attainment of orbital velocity in a re-entry capable single stage demands extraordinary propulsive performance. Most studies to date have focused on high pressure hydrogen/oxygen (H2/O2) rocket engines for the primary propulsion of such vehicles. However it is the authors opinion that despite recent advances in materials technology such an approach is not destined to succeed, due to the relatively low specific impulse of this type of propulsion. Airbreathing engines offer a possible route forward with their intrinsically higher specific impulse. However their low thrust/weight ratio, limited Mach number range and high dynamic pressure trajectory have in the past cancelled any theoretical advantage. By design review of the relevant characteristics of both rockets and airbreathing engines this paper sets out the rationale for the selection of deeply precooled hybrid airbreathing rocket engines for the main propulsion system of SSTO launchers as exemplified by the SKYLON vehicle [1]. 2. Propulsion Candidates This paper will only consider those engine types which would result in politically and environmentally acceptable vehicles. Therefore engines employing nuclear reactions (eg: onboard fission reactors or external nuclear pulse) and chemical engines with toxic exhausts (eg: fluorine/oxygen) will be excluded. The candidate engines can be split into two broad groups, namely pure rockets and engines with an airbreathing component. Since none of the airbreathers are capable of accelerating an SSTO vehicle all the way to orbital velocity, a practical vehicle will always have an onboard rocket engine to complete the ascent. Therefore the use of airbreathing has always been proposed within the context of improving the specific impulse of pure rocket propulsion during the initial lower Mach portion of the trajectory. Airbreathing engines have a much lower thrust/ weight ratio than rocket engines (à ¢Ã¢â‚¬ °Ã‹â€ 10%) which tends to offset the advantage of reduced fuel consumption. Therefore vehicles with airbreathing engines invariably have wings and employ a lifting trajectory in order to reduce the installed thrust requirement and hence the airbreathing engine mass penalty. The combination of wings and airbreathing engines then demands a low flat trajectory (compared to a ballistic rocket trajectory) in order to maximise the installed performance (i.e. (thrust-drag)/fuel flow). This high dynamic pressure trajectory gives rise to one of the drawbacks of an airbreathing approach since the airframe heating and loading are increased during the ascent which ultimately reflects in increased structure mass. However the absolute level of mass growth depends on the relative severity of the ascent as compared with reentry which in turn is mostly dependant on the type of airbreathing engine selected. An a dditional drawback to the low trajectory is increased drag losses particularly since the vehicle loiters longer in the lower atmosphere due to the lower acceleration, offset to some extent by the much reduced gravity loss during the rocket powered ascent. Importantly however, the addition of a set of wings brings more than just performance advantages to airbreathing vehicles. They also give considerably increased abort capability since a properly configured vehicle can remain in stable flight with up to half of its propulsion systems shutdown. Also during reentry the presence of wings reduces the ballistic coefficient thereby reducing the heating and hence thermal protection system mass, whilst simultaneously improving the vehicle lift/drag ratio permitting greater crossrange. The suitability of the following engines to the SSTO launcher role will be discussed since these are representative of the main types presently under study within various organisations world-wide: Liquid Hydrogen/Oxygen rockets Ramjets and Scramjets Turbojets/Turborockets and variants Liquid Air Cycle Engines (LACE) and Air Collection Engines (ACE) Precooled hybrid airbreathing rocket engines (RB545/SABRE) 3.Selection Criteria The selection of an optimum propulsion system involves an assessment of a number of interdependant factors which are listed below. The relative importance of these factors depends on the severity of the mission and the vehicle characteristics. Engine performance Useable Mach number and altitude range. Installed specific impulse. Installed thrust/weight. Performance sensitivity to component level efficiencies. Engine/Airframe integration Effect on airframe layout (Cg/Cp pitch trim structural efficiency). Effect of required engine trajectory (Q and heating) on airframe technology/materials. Technology level Materials/structures/aerothermodynamic and manufacturing technology. Development cost Engine scale and technology level. Complexity and power demand of ground test facilities. Necessity of an X plane research project to precede the main development program. 4.Hydrogen/Oxygen Rocket Engines Hydrogen/oxygen rocket engines achieve a very high thrust/weight ratio (60-80) but relatively low specific impulse (450-475 secs in vacuum) compared with conventional airbreathing engines. Due to the relatively large à ¢Ã‹â€ Ã¢â‚¬  V needed to reach low earth orbit (approx 9 km/s including gravity and drag losses) in relation to the engine exhaust velocity, SSTO rocket vehicles are characterised by very high mass ratios and low payload fractions. The H2/O2 propellant combination is invariably chosen for SSTO rockets due to its higher performance than other alternatives despite the structural penalties of employing a very low density cryogenic fuel. In order to maximise the specific impulse, high area ratio nozzles are required which inevitably leads to a high chamber pressure cycle in order to give a compact installation and reduce back pressure losses at low altitude. The need to minimise back pressure losses normally results in the selection of some form of altitude compensating nozzle since conventional bell nozzles have high divergence and overexpansion losses when running in a separated condition. The high thrust/weight and low specific impulse of H2/O2 rocket engines favours vertical takeoff wingless vehicles since the wing mass and drag penalty of a lifting trajectory results in a smaller payload than a steep ballistic climb out of the atmosphere. The ascent trajectory is therefore extremely benign (in terms of dynamic pressure and heating) with vehicle material selection determined by re-entry. Relative to airbreathing vehicles a pure rocket vehicle has a higher density (gross take off weight/volume) due to the reduced hydrogen consumption which has a favourable effect on the tankage and thermal protection system mass. In their favour rocket engines represent broadly known (current) technology, are ground testable in simple facilities, functional throughout the whole Mach number range and physically very compact resulting in good engine/airframe integration. Abort capability for an SSTO rocket vehicle would be achieved by arranging a high takeoff thrust/weight ratio (eg: 1.5) and a large number of engines (eg: 10) to permit shutdown of at least two whilst retaining overall vehicle control. From an operational standpoint SSTO rockets will be relatively noisy since the high takeoff mass and thrust/weight ratio results in an installed thrust level up to 10 times higher than a well designed airbreather. Reentry should be relatively straightforward providing the vehicle reenters base first with active cooling of the engine nozzles and the vehicle base. However the maximum lift/drag ratio in this attitude is relatively low (approx 0.25) limiting the maximum achievable crossrange to around 250 km. Having reached a low altitude some of the main engines would be restarted to control the subsonic descent before finally effecting a tailfirst landing on legs. Low crossrange is not a particular problem providing the vehicle operator has adequate time to wait for the orbital plane to cross the landing site. However in the case of a military or commercial operator this could pose a serious operational restriction and is consequently considered to be an undesirable characteristic for a new launch vehicle. In an attempt to increase the crossrange capability some designs attempt nosefirst re-entry of a blunt cone shaped vehicle or alternatively a blended wing/body configuration. This approach potentially increases the lift/drag ratio by reducing the fuselage wave drag and/or increasing the aerodynamic lift generation. However the drawback to this approach is that the nosefirst attitude is aerodynamically unstable since the aft mounted engine package pulls the empty center of gravity a considerable distance behind the hypersonic center of pressure. The resulting pitching moment is difficult to trim without adding nose ballast or large control surfaces projecting from the vehicle base. It is expected that the additional mass of these components is likely to erode the small payload capability of this engine/vehicle combination to the point where it is no longer feasible. Recent advances in materials technology (eg: fibre reinforced plastics and ceramics) have made a big impact on the feasibility of these vehicles. However the payload fraction is still very small at around 1-2% for an Equatorial low Earth orbit falling to as low as 0.25% for a Polar orbit. The low payload fraction is generally perceived to be the main disadvantage of this engine/vehicle combination and has historically prevented the development of such vehicles, since it is felt that a small degree of optimism in the preliminary mass estimates may be concealing the fact that the real payload fraction is negative. One possible route forward to increasing the average specific impulse of rocket vehicles is to employ the atmosphere for both oxidiser and reaction mass for part of the ascent. This is an old idea dating back to the 1950s and revitalised by the emergence of the BAe/Rolls Royce HOTOL project in the 1980s [2]. The following sections will review the main airbreathing engine candidates and trace the design background of precooled hybrid airbreathing rockets. 5.Ramjet and Scramjet Engines A ramjet engine is from a thermodynamic viewpoint a very simple device consisting of an intake, combustion and nozzle system in which the cycle pressure rise is achieved purely by ram compression. Consequently a separate propulsion system is needed to accelerate the vehicle to speeds at which the ramjet can takeover (Mach 1-2). A conventional hydrogen fuelled ramjet with a subsonic combustor is capable of operating up to around Mach 5-6 at which point the limiting effects of dissociation reduce the effective heat addition to the airflow resulting in a rapid loss in nett thrust. The idea behind the scramjet engine is to avoid the dissociation limit by only partially slowing the airstream through the intake system (thereby reducing the static temperature rise) and hence permitting greater useful heat addition in the now supersonic combustor. By this means scramjet engines offer the tantalising prospect of achieving a high specific impulse up to very high Mach numbers. The consequent de crease in the rocket powered à ¢Ã‹â€ Ã¢â‚¬  V would translate into a large saving in the mass of liquid oxygen required and hence possibly a reduction in launch mass. Although the scramjet is theoretically capable of generating positive nett thrust to a significant fraction of orbital velocity it is unworkable at low supersonic speeds. Therefore it is generally proposed that the internal geometry be reconfigured to function as a conventional ramjet to Mach 5 followed by transition to scramjet mode. A further reduction of the useful speed range of the scramjet results from consideration of the nett vehicle specific impulse ((thrust-drag)/fuel flow) in scramjet mode as compared with rocket mode. This tradeoff shows that it is more effective to shut the scramjet down at Mach 12-15 and continue the remainder of the ascent on pure rocket power. Therefore a scramjet powered launcher would have four main propulsion modes: a low speed accelerator mode to ramjet followed by scramjet and finally rocket mode. The proposed low speed propulsor is often a ducted ejector rocket system employing the scramjet injector struts as both ejector nozzles to entrain air at low speeds and later as the rocket combustion chambers for the final ascent. Whilst the scramjet engine is thermodynamically simple in conception, in engineering practice it is the most complex and technically demanding of all the engine concepts discussed in this paper. To make matters worse many studies including the recent ESA Winged Launcher Concept study have failed to show a positive payload for a scramjet powered SSTO since the fundamental propulsive characteristics of scramjets are poorly suited to the launcher role. The low specific thrust and high specific impulse of scramjets tends to favour a cruise vehicle application flying at fixed Mach number over long distances, especially since this would enable the elimination of most of the variable geometry. Scramjet engines have a relatively low specific thrust (nett thrust/airflow) due to the moderate combustor temperature rise and pressure ratio, and therefore a very large air mass flow is required to give adequate vehicle thrust/weight ratio. However at constant freestream dynamic head the captured air mass flow reduces for a given intake area as speed rises above Mach 1. Consequently the entire vehicle frontal area is needed to serve as an intake at scramjet speeds and similarly the exhaust flow has to be re-expanded back into the original streamtube in order to achieve a reasonable exhaust velocity. However employing the vehicle forebody and aftbody as part of the propulsion system has many disadvantages: The forebody boundary layer (up to 40% of the intake flow) must be carried through the entire shock system with consequent likelihood of upsetting the intake flow stability. The conventional solution of bleeding the boundary layer off would be unacceptable due to the prohibitive momentum drag penalty. The vehicle undersurface must be flat in order to provide a reasonably uniform flowfield for the engine installation. The flattened vehicle cross section is poorly suited to pressurised tankage and has a higher surface area/volume than a circular cross section with knock-on penalties in aeroshell, insulation and structure mass. Since the engine and airframe are physically inseparable little freedom is available to the designer to control the vehicle pitch balance. The single sided intake and nozzle systems positioned underneath the vehicle generate both lift and pitching moments. Since it is necessary to optimise the intake and nozzle system geometry to maximise the engine performance it is extremely unlikely that the vehicle will be pitch balanced over the entire Mach number range. Further it is not clear whether adequate CG movement to trim the vehicle could be achieved by active propellant transfer. Clustering the engines into a compact package underneath the vehicle results in a highly interdependant flowfield. An unexpected failure in one engine with a consequent loss of internal flow is likely to unstart the entire engine installation precipitating a violent change in vehicle pitching moment. In order to focus the intake shock system and generate the correct duct flow areas over the whole Mach range, variable geometry intake/combustor and nozzle surfaces are required. The large variation in flow passage shape forces the adoption of a rectangular engine cross section with flat moving ramps thereby incurring a severe penalty in the pressure vessel mass. Also to maximise the installed engine performance requires a high dynamic pressure trajectory which in combination with the high Mach number imposes severe heating rates on the airframe. Active cooling of significant portions of the airframe will be necessary with further penalties in mass and complexity. Further drawbacks to the scramjet concept are evident in many areas. The nett thrust of a scramjet engine is very sensitive to the intake, combustion and nozzle efficiencies due to the exceptionally poor work ratio of the cycle. Since the exhaust velocity is only slightly greater than the incoming freestream velocity a small reduction in pressure recovery or combustion efficiency is likely to convert a small nett thrust into a small nett drag. This situation might be tolerable if the theoretical methods (CFD codes) and engineering knowledge were on a very solid footing with ample correlation of theory with experiment. However the reality is that the component efficiencies are dependant on the detailed physics of poorly understood areas like flow turbulence, shock wave/boundary layer interactions and boundary layer transition. To exacerbate this deficiency in the underlying physics existing ground test facilities are unable to replicate the flowfield at physically representative sizes , forcing the adoption of expensive flight research vehicles to acquire the necessary data. Scramjet development could only proceed after a lengthy technology program and even then would probably be a risky and expensive project. In 1993 Reaction Engines estimated that a 130 tonne scramjet vehicle development program would cost $25B (at fixed prices) assuming that the program proceeded according to plan. This program would have included two X planes, one devoted to the subsonic handling and low supersonic regime and the other an air dropped scramjet research vehicle to explore the Mach 5-15 regime. 6.Turbojets, Turborockets and Variants In this section are grouped those engines that employ turbocompressors to compress the airflow but without the aid of precoolers. The advantage of cycles that employ onboard work transfer to the airflow is that they are capable of operation from sea level static conditions. This has important performance advantages over engines employing solely ram compression and additionally enables a cheaper development program since the mechanical reliability can be acquired in relatively inexpensive open air ground test facilities. 6.1 Turbojets Turbojets (Fig. 1) exhibit a very rapid thrust decay above about Mach 3 due to the effects of the rising compressor inlet temperature forcing a reduction in both flow and pressure ratio. Compressors must be operated within a stable part of their characteristic bounded by the surge and choke limits. In addition structural considerations impose an upper outlet temperature and spool speed limit. As inlet temperature rises (whilst operating at constant Wà ¢Ã‹â€ Ã… ¡T/P and N/à ¢Ã‹â€ Ã… ¡T) the spool speed and/or outlet temperature limit is rapidly approached. Either way it is necessary to throttle the engine by moving down the running line, in the process reducing both flow and pressure ratio. The consequent reduction in nozzle pressure ratio and mass flow results in a rapid loss in nett thrust. However at Mach 3 the vehicle has received an insufficient boost to make up for the mass penalty of the airbreathing engine. Therefore all these cycles tend to be proposed in conjunction with a subsonic combustion ramjet mode to higher Mach numbers. The turbojet would be isolated from the hot airflow in ramjet mode by blocker doors which allow the airstream to flow around the core engine with small pressure loss. The ramjet mode provides reasonable specific thrust to around Mach 6-7 at which point transition to rocket propulsion is effected. Despite the ramjet extension to the Mach number range the performance of these systems is poor due mainly to their low thrust/weight ratio. An uninstalled turbojet has a thrust/weight ratio of around 10. However this falls to 5 or less when the intake and nozzle systems are added which compares badly with a H2/O2 rocket of 60+. 6.2 Turborocket The turborocket (Fig. 2) cycles represent an attempt to improve on the low thrust/weight of the turbojet and to increase the useful Mach number range. The pure turborocket consists of a low pressure ratio fan driven by an entirely separate turbine employing H2/O2 combustion products. Due to the separate turbine working fluid the matching problems of the turbojet are eased since the compressor can in principle be operated anywhere on its characteristic. By manufacturing the compressor components in a suitable high temperature material (such as reinforced ceramic) it is possible to eliminate the ramjet bypass duct and operate the engine to Mach 5-6 whilst staying within outlet temperature and spool speed limits. In practice this involves operating at reduced nondimensional speed N/à ¢Ã‹â€ Ã… ¡T and hence pressure ratio. Consequently to avoid choking the compressor outlet guide vanes a low pressure ratio compressor is selected (often only 2 stages) which permits operation over a wider flow range. The turborocket is considerably lighter than a turbojet. However the low cycle pressure ratio reduces the specific thrust at low Mach numbers and in conjunction with the preburner liquid oxygen flow results in a poor specific impulse compared to the turbojet. 6.3 Expander Cycle Turborocket This cycle is a variant of the turborocket whereby the turbine working fluid is replaced by high pressure regeneratively heated hydrogen warmed in a heat exchanger located in the exhaust duct (Fig. 3). Due to heat exchanger metal temperature limitations the combustion process is normally split into two stages (upstream and downstream of the ma- LHLH LOx/LH2 Fig. 1 Turbo-ramjet Engine (with integrated rocket engine). LOx/LH2LH2 LOx/LH2 Fig. 2 Turborocket. LH2LOx/LH2 Fig. 3 Turbo-expander engine. trix) and the turbine entry temperature is quite low at around 950K. This variant exhibits a moderate improvement in specific impulse compared with the pure turborocket due to the elimination of the liquid oxygen flow. However this is achieved at the expense of additional pressure loss in the air ducting and the mass penalty of the heat exchanger. Unfortunately none of the above engines exhibit any performance improvement over a pure rocket approach to the SSTO launcher problem, despite the wide variations in core engine cycle and machinery. This is for the simple reason that the core engine masses are swamped by the much larger masses of the intake and nozzle systems which tend to outweigh the advantage of increased specific impulse. Due to the relatively low pressure ratio ramjet modes of these engines, it is essential to provide an efficient high pressure recovery variable geometry intake and a variable geometry exhaust nozzle. The need for high pressure recovery forces the adoption of 2 dimensional geometry for the intake system due to the requirement to focus multiple oblique shockwaves over a wide mach number range. This results in a very serious mass penalty due to the inefficient pressure vessel cross section and the physically large and complicated moving ramp assembly with its high actuation loads. Similarly the exhaust nozzle geometry must be capable of a wide area ratio variation in order to cope with the widely differing flow conditions (Wà ¢Ã‹â€ Ã… ¡T/P and pressure ratio) between transonic and high Mach number flight. A further complication emerges due to the requirement to integrate the rocket engine needed for the later ascent into the airbreathing engine nozzle. This avoids the prohibitive base drag penalty that would result from a separate dead nozzle system as the vehicle attempted to accelerate through transonic. 7. Liquid Air Cycle Engines (LACE) and Air Collection Engines (ACE) Liquid Air Cycle Engines were first proposed by Marquardt in the early 1960s. The simple LACE engine exploits the low temperature and high specific heat of liquid hydrogen in order to liquify the captured airstream in a specially designed condenser (Fig. 4). Following liquifaction the air is relatively easily pumped up to such high pressures that it can be fed into a conventional rocket combustion chamber. The main advantage of this approach is that the airbreathing and rocket propulsion systems can be combined with only a single nozzle required for both modes. This results in a mass saving and a compact installation with efficient base area utilisation. Also the engine is in principle capable of operation from sea level static conditions up to perhaps Mach 6-7. LH2 LO2 Liquid Air Turbopump Fig. 4 Liquid Air Cycle Engine (LACE). The main disadvantage of the LACE engine however is that the fuel consumption is very high (compared to other airbreathing engines) with a specific impulse of only about 800 secs. Condensing the airflow necessitates the removal of the latent heat of vaporisation under isothermal conditions. However the hydrogen coolant is in a supercritical state following compression in the turbopump and absorbs the heat load with an accompanying increase in temperature. Consequently a temperature pinch point occurs within the condenser at around 80K and can only be controlled by increasing the hydrogen flow to several times stoichiometric. The air pressure within the condenser affects the latent heat of vaporisation and the liquifaction temperature and consequently has a strong effect on the fuel/air ratio. However at sea level static conditions of around 1 bar the minimum fuel/air ratio required is about 0.35 (ie: 12 times greater than the stoichiometric ratio of 0.029) assuming that the hydrogen had been compressed to 200 bar. Increasing the air pressure or reducing the hydrogen pump delivery pressure (and temperature) could reduce the fuel/ air ratio to perhaps 0.2 but nevertheless the fuel flow remains very high. At high Mach numbers the fuel flow may need to be increased further, due to heat exchanger metal temperature limitations (exacerbated by hydrogen embrittlement limiting the choice of tube materials). To reduce the fuel flow it is sometimes proposed to employ slush hydrogen and recirculate a portion of the coolant flow back into the tankage. However the handling of slush hydrogen poses difficult technical and operational problems. From a technology standpoint the main challenges of the simple LACE engine are the need to prevent clogging of the condenser by frozen carbon dioxide, argon and water vapour. Also the ability of the condenser to cope with a changing g vector and of designing a scavenge pump to operate with a very low NPSH inlet. Nevertheless performance studies of SSTOs equipped with LACE engines have shown no performance gains due to the inadequate specific impulse in airbreathing mode despite the reasonable thrust/weight ratio and Mach number capability. The Air Collection Engine (ACE) is a more complex variant of the LACE engine in which a liquid oxygen separator is incorporated after the air liquifier. The intention is to takeoff with the main liquid oxygen tanks empty and fill them during the airbreathing ascent thereby possibly reducing the undercarriage mass and installed thrust level. The ACE principal is often proposed for parallel operation with a ramjet main propulsion system. In this variant the hydrogen fuel flow would condense a quantity of air from which the oxygen would be separated before entering the ramjet combustion chamber at a near stoichiometric mixture ratio. The liquid nitrogen from the separator could perform various cooling duties before being fed back into the ramjet airflow to recover the momentum drag. The oxygen separator would be a complex and heavy item since the physical properties of liquid oxygen and nitrogen are very similar. However setting aside the engineering details, the basic thermodynamics of the ACE principal are wholly unsuited to an SSTO launcher. Since a fuel/air mixture ratio of approximately 0.2 is needed to liquify the air and since oxygen is 23.1% of the airflow it is apparent that a roughly equal mass of hydrogen is required to liquify a given mass of oxygen. Therefore there is no saving in the takeoff propellant loading and in reality a severe structure mass penalty due to the increased fuselage volume needed to contain the low density liquid hydrogen. 8. Precooled Hybrid Airbreathing Rocket Engines This last class of engines is specifically formulated for the SSTO propulsion role and combines some of the best features of the previous types whilst simultaneously overcoming their faults. The first engine of this type was the RB545 powerpla

Frederick Douglass, an American Slave Essay -- Autobiography, Douglas

The Narrative of the Life of Frederick Douglass, an American Slave, was the first of the three autobiographies that Frederick Douglass wrote himself. It’s a story about slavery and the meaning of freedom of the antebellum America. According to The Free Dictionary, Slavery is defined as the state or condition of being a slave; a civil relationship whereby one person has absolute power over another and controls his life, liberty, and fortune (freedictionary.com). Frederick Douglass’s book is about a bondage he obtained since birth; a slave for life. He was separated from his mother, Harriet Bailey, at birth and knew his father was white male. He lived on the â€Å"Great House Farm† plantation for his younger years; this is where he saw his first violent act towards a slave. Douglass went through many ups and downs. At the age of seven, he was moved to another house where he first learned reading and writing. However, He was beaten brutally so he can be â€Å"broke n† into a good disciplined slave. Douglass describes many elements in his narrative; Douglass explains how slaveholders were able to sustain themselves with their actions. Frederick describes the ways the slaves stayed where they were and did not attempt to escape. He also addresses a number of myths created by slaves and slaveholders that he wishes to prove wrong. In the Narrative of the Life of Frederick Douglass, an American Slave, Frederick Douglass describes the ways a slaveholder sustain their actions, ways a slave was kept from escaping and proves the myths of slaves and slaveholders wrong. Slaveholders had a number of ways to justify themselves for their actions according to Douglass. One way they justify themselves for their actions was that slaves were lower than animals.... ... anguish (19).† In all, south was far what the images of fancy and big, yet depressing and unrealistic. All in all, Frederick Douglass’s book, Narrative of the Life of Frederick Douglass, an American Slave, was a story of slavery and freedom. He was fortunate that he was able to experience a better slave life than others. He was able to obtain knowledge about reading that he was not obtaining to be a slave for all his life. He, unlike other slaves, knew he was not supposed to be a slave for the rest of his life. He described the ways by which slaveholders justify themselves for their actions. He was one of the rare ones who did not lose their way to freedom; he discussed the many ways that slaves were kept from thinking about escaping and freedom. Once he was free, he wrote this Narrative and refutes many myths that many have said about slaves and slaveholders.